46th International Conference on Environmental Systems ICES-2016-119 10-14 July 2016, Vienna, Austria

Versatile Thermal Insulation for Cryogenic Upper Stages

M. Moser1 and W. Hoidn2 RUAG Space GmbH, Stachegasse 16, 1120 Vienna, Austria

P. Delouard3 RUAG Space AG, Schaffhauserstrasse 580, 8052 Zürich, Switzerland

and

A. Tvaruzka4, M. Loche5 and U. Lafont6 European Space Agency ESA, Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands

Within ESAs future launcher preparatory program RUAG Space has developed a Versatile Thermal Insulation (VTI) system for cryogenic launcher upper-stages. The pneumatically deployable VTI system is self-sustained and mounted via interchangeable segments on the payload interface adapter. At this location the system is protected from environments on the launch pad and aerothermal loads during lift-off. On the interface adapter the VTI is independent from the upper stage, which in turn does not need to be modified for the insulation system. As the insulating VTI segments are adaptable in size the system can be designed to any cryogenic upper stage configuration. VTI is thus a viable upgrade kit for existing- and future launchers.

In orbit VTI segments deploy on command through pressurization in a sequence that concludes in an insulating skirt covering the cryogenic stage. The entire deployment process is only controlled by pressure and built-in retainers made of hook/loop fasteners. Apart from the hold-down and release assembly no additional mechanism is needed for deployment. This approach reduces system complexity and leads to a mass below 80kg for a 75m2 covering VTI system.

The deployment process has been successfully verified with a full scale, 1.5 meter by 5 meter, demonstrator model of one VTI segment. It was demonstrated that deployment is reliable and reproducible in real size and that stowing in rolled form is compatible with launch loads. Thermal analyses have shown that on a generic 4.4m upper stage the VTI system reduces heat input into the liquid hydrogen tank by more than 90% resulting in a strongly reduced fuel boil-off rate.

Nomenclature AIT = Assembly, Integration and Test CAD = Computer Aided Design CFRP = Carbon Fiber Reinforced Plastic

1 Lead Systems Engineer VTI, Thermal Systems, RUAG Space GmbH, Stachegasse 16, A-1120 Vienna, Austria 2 Senior Systems Engineer, Thermal Systems, RUAG Space GmbH, Stachegasse 16, A-1120 Vienna, Austria 3 Senior Thermal Engineer, Launchers, RUAG Space AG, Schaffhauserstrasse 580, CH-8052 Zürich, Switzerland. 4 Mechanisms Engineer, TEC-MSM, Mechanical Engineering Department, ESA/ESTEC, Keplerlaan 1, NL-2201AZ Noordwijk, The Netherlands. 5 Thermal Engineer, TEC-MTT, Mechanical Engineering Department, ESA/ESTEC, Keplerlaan 1, NL-2201AZ Noordwijk, The Netherlands. 6 Materials Engineer, TEC-QTE, Product Assurance & Safety Department, ESA/ESTEC, Keplerlaan 1, NL-2201AZ Noordwijk, The Netherlands. CUS = Cryogenic Upper Stage ESA = European Space Agency FLPP = Future Launcher Preparatory Program GEO = Geostationary Orbit GH2 = Gaseous Hydrogen GHe = Gaseous Helium GN2 = Gaseous Nitrogen GMM = Geometrical Mathematical Model GSE = Ground Support Equipment GTO = Geostationary Tranfer Orbit HDRM = Hold Down and Release Mechanism LH2 = Liquid Hydrogen LOx = Liquid Oxygen MLI = Multilayer Insulation MMOD = Micro Meteorite or Orbital Debris PET = Polyethylene terephthalate SLI = Single Layer Insulation TMM = Thermal Mathematical Model TRL = Technology Readiness Level ULA = United Launch Alliance VDA = Vacuum Deposited Aluminum VTI = Versatile Thermal Insulation

I. Introduction EDUCTION of propellant boil-off allows cryogenic upper stages to perform longer coasting phases with R multiple engine re-ignitions and thus improvement of overall launcher performance and payload orbital insertion flexibility. MLI has been shown to be an excellent barrier against incident thermal energy and in consequence an extremely effective thermal insulation system on cryogenic tanks. Several systems have been developed in recent years to further improve MLI performance at cryogenic temperatures allowing for increased- or even long duration storage of cryogenic propellants in orbit.1-4 However, MLI cannot be built robust enough to survive aerodynamic loads appearing during typical rocket launches. As a consequence it is common practice that thick layers of foam insulation are sprayed onto the external walls of launch vehicle cryogenic stages. These foams offer good environmental protection on ground but worse thermal performance than MLI in orbit and also tend to degrade upon launch vehicle ascent resulting in high fuel boil-off rates and strong restrictions in cryogenic upper stage (CUS) mission duration and utilisation flexibility. It is thus from highest interest to develop a system that supports increasing CUS performance by maintaining its properties during launch and ascent and in addition features the excellent insulation properties of MLI in orbit. As this increased performance is not needed on all missions, designing a stage to always comply with these long coasting or multi-ignition mission scenarios would result in reduced overall launcher performance due to increased mass penalty. Consequently, the main goal of the FLPP VTI project was the development of an add-on CUS insulation system which is to be applied only in cases required. The insulation shall be versatile in terms of applicability, adaptability to stage- or payload configuration and AIT constraints. The overall footprint i.e. the number of interfacing areas on the stage shall be minimal. With its expertise in developing advanced insulation products RUAG Space has been involved in studying VTI concepts since 2009. Based on a detailed requirements definition with industry three general approaches were identified for further investigation:  Aerothermal Shroud  Mechanically Deployable Cover  Pneumatically Deployable System In course of a trade-off process, both, the Aerothermal Shroud and Mechanically Deployable Cover were discarded as they revealed major technical challenges and performance drawbacks. The former, a cylindrical shroud which covers MLI during launch, was shown to have high mass and that its thermal performance benefit largely depends on a compromise between mechanical stability and attachment design. The Mechanically Deployable Cover

2 International Conference on Environmental Systems

would deploy in orbit but with the help of mechanisms and mechanical actuators. This deployment approach would thus result in large system complexity, and depending on the deployment method chosen, high mass penalty. Further, both systems require a number of interfaces which permanently have to be foreseen on the upper stage. The Pneumatically Deployable System on the other hand offers a number of advantages:  it can be built with a low number of mechanisms and mechanical components  it is of lightest weight of the investigated systems  it can be stowed within a small volume. In combination, this results in a system with low complexity in transport, handling and integration and, based on the exceptional packing efficiency, provides the stage architect with a high degree of freedom for placement of the system on the CUS.5 These facts have also been recognized by ULA who have developed a sun shield for an Atlas CUS in cooperation with ILC Dover.4 There straight beams are stowed in an accordion pattern and pneumatically deployed using rigid end-fittings. Based on the above trade-off we have focused on the Pneumatically Deployable VTI System, which has been designed and developed in the frame of an ESA FLPP study since 2012.

II. The VTI System

A. Overview The Pneumatically Deployable VTI System was developed for a generic cryogenic upper stage with ø4.4m diameter and LH2 and LOx tanks. VTI was required to only cover the hydrogen tank. VTI consists of several segments of a deployable insulation which are attached on the external side of the payload interface adapter Versatile Thermal Insulation between payload and upper-stage cryogenic deployed tanks. For on-ground-operations, launch and ascent the system is protected against environmental hazards and aerothermal loads by encapsulation under the fairing. Upon fairing jettison and when the CUS has reached a stable coasting phase VTI is deployed pneumatically over the LH2 tank and Payload is, in its baseline configuration, designed to keep full functionality for up to 10 hours. With Cryogenic Stage further upgrades the system may however be designed for longer mission durations and utilized for solar sails, orbital propellant depots Figure 1. Illustration of VTI in Orbit or telescope sun shields. A principle illustration of the deployed VTI system covering the generic upper stage LH2 tank is presented in Figure 1. For the ø4.4m stage configuration 10 VTI segments, each 1.5m wide, are circumferentially attached to the payload adapter and are divided in groups of 5 so-called “inner” segments and 5 ”outer” segments. Outer segments, are mounted to the payload adapter an elevated (+Z, i.e. flight direction) position with respect to the inner segments. The VTI assembly in stowed configuration, mounted on the payload adapter, is shown in the CAD image Figure 2.

Designed with a modular layout, each segment is self-contained, interchangeable and is composed of two major sub-systems: 1. Storage and Deployment Support System, which itself is composed of a. Storage and Release Subsystem it retains a VTI segment during launch and early orbital phases in rolled form of app. 200mm diameter. The Hold Down and Release Mechanism (HDRM) subsystem consists of a spring-load driven actuator that locks retaining straps and a protective cover foil until the release signal is given. b. Pneumatic Subsystem Interfaces consisting of valves, pressure transducers, and nozzles required to feed pressurizing gases into the inflatable support structure of each segment. 3 International Conference on Environmental Systems

c. Mechanical Support Structure is the structural interface between payload adapter and VTI and forms the backbone of each segment. Both, the storage and release subsystem and the pneumatic subsystem interfaces are mounted to the support structure. It further provides the fixation elements for the components of the Deployable Insulation. 2. Deployable Insulation, composed of a. Insulation Blankets of space facing, 10 layer MLI blankets with Kapton®/VDA outer layer and internal layers of lightweight 3µm VDA/PET/VDA foils separated by non-woven spacer fleeces. A double sided VDA coated Mylar® SLI is facing towards the cryogenic LH2 tank. MLI and SLI are held at distance by the inflatable support structure, providing effective, radiatively decoupled layers. b. Inflatable Support Structure consisting of two inflatable L-shaped beams which incorporate the interfaces to the Mechanical Support Structure. The beams are interconnected with CFRP rods and conically shaped with a diameter of 100mm towards the payload adapter and 50mm at the end of the deployed system. This conical shape results in a reduction of needed gas volume, reduced VTI distance from the tank and improved alignment along the CUS. In the baseline VTI configuration the beams are made of double sided VDA coated Mylar® which is formed into a tubular shaped, gastight bladder.

Outer Segments (5x) Common Subsystem allowable Area

Inner Segments Payload Interface (5x) Adapter

Figure 2. VTI segments in stowed configuration on payload adapter

Pneumatic, power and command subsystems are common to the VTI assembly and have interfaces to all segments. These subsystems are directly mounted to the payload adapter in the allowable area indicated in Figure 2 between VTI segments and payload. They will be covered with MLI and actively thermal-controlled. Within the pneumatic subsystem the pressurization gas is stored and distributed to the individual VTI segments. The layout has analogies to pressure control assemblies of bi-propellant satellites, which allows designing this subsystem entirely of spaceflight heritage components. In the baseline VTI configuration the system is built self- sustained from the upper stage and GN2, supplied from distinct containers, is used to pressurize the deployable segments. In dependence of the launcher configuration also GHe or GH2 boil-off from the cryogenic tank could be employed for VTI pressurization. The total mass of a VTI system as described above, designed to cover the LH2 tank of a ø4.4m CUS, is assessed with less than 80kg.

B. Deployment Sequence The deployment sequence of the VTI segments, presented in yellow color in Figure 3, starts from the stowed VTI configuration (0), where the deployable insulation is retained in rolled form on the payload adapter and protected from environmental hazards by a protective cover cloth.

4 International Conference on Environmental Systems

When the CUS reaches a stable coasting phase a signal is given which triggers release and deployment of the VTI segments in two groups of five inner and five outer segments, respectively.

Figure 3. VTI deployment sequence

First the five inner segments are deployed - steps (1) to (4) in Figure 3. Upon confirmation of opening of the hold-down device the deployable insulation is pressurized with gaseous nitrogen. Pressure build-up in the inflatable beams of a segment, triggers the insulation to deploy in an uncoiling motion. Using hook/loop fasteners of defined type and dimension this uncoiling motion is controlled in such a way that first segments deploy horizontally, in a safe corridor perpendicular to the main stage axis (Z-axis), until the entire insulation is unrolled - see Figure 3, (1) to (2). A comparable method using Velcros® to control uncurling motion of inflating beams has been described in literature and was reported to show excellent reproducibility and controllability when deployed in a 0g environment.5, 6 Further pressure build-up triggers a specially designed hook/loop fastener retained part of the inflatable truss, the bending element, to unfold resulting in a bending motion towards the stage main axis along the cryogenic tank (3). The bending motion is actuated only by the pressure within the tubes and performed passively, without the need of an externally controlled mechanical system. Thus the fully deployed inner segments partially cover the stage without contact to the tank (4) at a distance defined by the L-shaped profile of the fully pressurized inflatable support structure. This deployment sequence with hook/loop fastener controlled uncoiling and pressure triggered bending motion was successfully and repeatedly demonstrated on a full scale model of one segment of 5m length and 1.5m width using off-loading balloons to simulate a 0g environment. The full demonstration test campaign is described in detail in paragraph IV below. With the inner segments folded towards the LH2 tank the first part of the deployment sequence is completed (4) and deployment of the five outer segments is started. Again these outer segments are first deployed horizontally until the entire insulation is unrolled as indicated in Figure 3, (5). Then the segments are bent (6), as described above, towards the tank to an end-positon on the inner segments resulting in the VTI fully deployed configuration (7). Outer segments overlap the inner segments by app. 100mm on each side resulting in reduced heat input towards the cryogenic stage. Hook/loop fasteners used for controlled uncoiling will also be employed for attaching outer to inner segments with an outer segment having a loop fastener oriented towards the tank and the inner segments having a hook fasteners orientated towards the bending outer segments. In that way a circumferentially closed cylindrical insulation system is achieved increasing mechanical robustness and thermal form fit. When the VTI system is fully deployed, pressure in the inflatable beams is reduced to minimum levels required to keep beam stiffness and avoid buckling upon stage maneuvers or engine reignition. This reduction in pressure levels reduces stresses to the inflatable bladder and improves the consumable gas budget. In case of pressure loss due to leakage or MMOD strike the affected segment is taken off the pressure feed. Still, due to the above described interconnection to neighboring segments VTI shape keeping is guaranteed in such an event. Nonetheless, leakage upon VTI deployment needs to be avoided as it might result in full loss of the respective segment.

5 International Conference on Environmental Systems

In comparison to existing technologies like Ref. 4 the herein presented system offers numerous advantages. While both overlapping of insulating segments and an alignment closer to the CUS improve the thermal performance of the system, the closer shape alignment, resulting from the bending motion of the segments, allows orienting the stage in a wider corridor without risking undesirable suntrapping or engine heat input. Implementation of further bending elements would even allow to fully shield the cryogenic tanks against external heat input. Using only HDRMs but no other mechanisms to control deployment, results in a low weight system with high design flexibility. Based on these advantages compared to state-of the art technology, the RUAG VTI system of a pneumatically deployable thermal insulation for cryogenic tanks has been filed for European patent application.7

C. Operation In Figure 4 the VTI mission timeline with respect to launcher upper stage operations is presented. Upon launch and early orbital phases VTI segments are stowed by their respective storage and release systems as shown in Figure 3, (1). After fairing jettison and a >10 minutes ignition of the upper stage engine (stage boost #1) the stage enters a stable orbital coasting phase.

4. Stage Ballistic Phase VTI deployment 6. Stage re-ignition 7. Payload separation VTI deployed VTI passivation 3. Stage boost #1

5. Stage Ballistic Flight 2. Fairing Jettison VTI deployed & functional

1. Launch

Mission Duration 30min 10 hours

Figure 4. VTI mission timeline

To avoid trapping of upper stage intrinsic heat from the launch loads in a deployed VTI system a stage chill down is performed before start of VTI segment deployment. With a command signal given either by stage or directly the VTI command & control unit the VTI starts the two staged deployment phase. Based on the full-scale demonstrator model (paragraph IV) the entire deployment sequence of 10 segments is timed to take less than 5 minutes for the 75m2 baseline system. When all segments are deployed covering the cryogenic tank the VTI enters the operational phase of about 10 hours. Within this phase the CUS may perform a series of maneuvers with short duration engine boosts or payload injections. The VTI segments will flex upon stage maneuver accelerations but return in their original shape providing full functionality during this period. Pressure is maintained and regulated in the deployable elements. The operational phase of the VTI ends with injection of the last payload and the maneuvers performed by the CUS to enter the passivation and possible deorbit period. During this phase also the VTI is passivated and residual pressurizing agents are vented in accordance with ESA clean-space regulations.8 After passivation the unpressurized VTI segments remain attached to the stage.

6 International Conference on Environmental Systems

III. Thermal Analyses

A. Analyses and Cases Based on the VTI CAD design a Geometrical Mathematical Model (GMM) and Thermal Mathematical Model (TMM) were established for thermal analyses. The VTI TMM is a middle size model with about 5800 thermal nodes and a similar number of surfaces in the GMMs. Nominal and off-nominal cases were calculated transient, performance and temperature distributions were derived versus mission duration. The nominal cases investigated included a launch case to assess the pre-release temperatures of the VTI sub- systems and provide realistic initial temperatures for the in-orbit cases. As a baseline for the analyses GEO missions were assumed. All in-orbit cases are calculated over half a GTO orbit from perigee to apogee and comprise a propulsion phase at the start and a two hour eclipse phase at the end. This approach made it possible to cover in a single case the worst hot (start, engine ignition) and cold (end of eclipse) situations. The temperatures at the end of hot launch case were set as the start temperatures of orbital cases. From all possible Sun orientation angles (angle between Sun vector and Launcher axis), four potential worst GTO cases as illustrated in Figure 5 have been selected for in-depth analyses: 90°  0°: glancing angle Sun illumination on tank 45° and tank covering VTI, maximum Sun illumination on VTI parts perpendicular to stage main axis 180°  45°: maximum Sun illumination on VTI subsystems 0°  90°: maximum Sun input and potential input through segment overlaps Figure 5. Sun-angle orientation

 180°: maximum Sun trapping case All cases consider a constant stage rotation at 0.2° per second, are transient, at winter solstice (for maximum Sun flux) and with pristine thermo-optical properties of VTI insulation materials. Heat flows and transient temperatures were calculated in all cases for an uncovered stage and compared with VTI in place. Results of the analyses are provided in Table 1, while the Figures below give examples of the analysis results derived.

B. Thermal Analyses Results In Figure 6 and Figure 7 below, radiative heat flows through the VTI segments and temperature on the MLI outer layer are presented for the nominal case with 90° solar illumination, repectively. Heat flows shown in Figure 6, are typically comprised between -10 and +7 W in dependence of segment rotatory

Figure 6. Heat Flow through individual VTI segments – 90° solar illumination 7 International Conference on Environmental Systems

orientation with respect to the sun. Flows are somewhat larger for the outer segments than for the inner ones. The later receive less solar input as they are partially overlapped by the outer segments and the 50mm – 100mm wide segment edges are not exposed in contrast to the outer segments. The corresponding average heat flows are 1.5 to 2W for the outer segments and less then 0.5 W for the inner ones. With the Kapton®/VDA MLI outer layer external surfaces of the VTI segments show temperature variations of up to 75°C during one rotation. Temperatures on the outer segments’ space facing layers vary between -85°C to -10°C, while variation on the inner segments is smaller approximately between -85°C and -20°C on segment center areas. As an effect of overlapping temperature variations on the inner segments edges are calculated to only between -50°C and -30°C. Temperatures along the segments length were found relatively constant limiting effects of thermal distortion on the inflatable beams. When the upper stage with VTI enters shadow, the temperature of the segment drops to less than -100°C in most areas. Still temperature distribution analyses confirmed that GN2 can safely be used for segment deployment. Components on the payload adapter will need to be wrapped in MLI blankets to keep them in safe operation margins.

Figure 7. Temperatures on VTI space facing surfaces - 90° solar illumination

The main results of thermal analyses, the heat flows into the tank, are presented in Table 1. For cases where the Sun is directly illuminating the stage i.e. 45°, 90°, the VTI system considerably reduces heat input into the LH2 tanks by more than a factor of 10. Both cases with glancing solar illumination, i.e. the 0° and the 180° case result in increases in the absorbed power on the tank, the latter due to sun-trapping between tank and deployed VTI system. Consequently these cases should be avoided when the VTI is used and considered in defining the proper stage trajectories. Table 1. Power absorbed by the LH2 tank (W/m2) before eclipse Analyses Case Bare Stage VTI 0° 0.60 5.55 45° 102.29 8.52 90° 129.11 10.26 180° 0.45 10.10

Sensitivity analyse have shown that further improvements in VTI performance can be achieved with higher stage rotation velocity and, most importantly, by avoiding VTI deployment directly after CUS boost#1 when the stage is still hot from launch loads and engine operation. This stage intrinsic heat would be trapped by the deployed VTI

8 International Conference on Environmental Systems

even increasing fuel boil-off in the early stages of the mission. A stage chill down period is thus recommended before start of VTI operation. To evaluate the failure effects and to quantify possible system degradation off-nominal cases were analysed where one and five VTI, i.e. all inner segments, were assumed to not deploy. As expected, failure of segments results in a considerable increase in absorbed heat by the tank. When one segment is not deployed average power absorbed is increased by 13 W/m2, with the failure of 5 segments an increase of 58 W/m2 has to be considered. All off-nominal cases were done for 90° solar illumination and thus represent a worst case scenario.

IV. Full-Scale Demonstrator Testing

A. Test flow and Setup A test flow was chosen where a demonstrator, representing one full-size VTI segment, is deployed prior and post vibration testing. For later utilization of the system, flight items will have to undergo a comparable sequence of 1. an initial deployment on ground to check functionality and perform final inspection, 2. followed by a launch campaign and mechanical exposure due to the launch environment 3. the in-orbit deployment of the VTI system – i.e. a second deployment. Thus the selected test flow is representative for the typical life cycle of a VTI segment. All (a) (b) tests were recorded with videos and photographic images and concluded by a detailed visual inspection of the demonstrator test items. For the deployment tests of the full-size demonstrator segment a GSE trolley was designed. It served as a supporting element for the vertically mounted inflatable structure. Additionally the GSE trolley provided the pneumatic control unit to regulate gas supply. The pneumatic demonstrator subsystem was designed to process Helium, Nitrogen and pressurized air. As a consequence of its on-site (c) availability for the deployment tests pressurized air was used. Gravity compensation was achieved using Helium balloons on suspension strings that were attached at two positions of the deployable insulation. The strings were equipped with swivels on both ends to provide additional degrees of freedom upon rotatory deployment of the inflating segment.

B. Demonstrator Deployment (d) The deployment sequence of the demonstrator test segment is presented in Figure 8. Image Figure 8 - (a) shows the GSE framework structure with the 1.5m wide test item mounted vertically. Suspension strings of 1,5m the gravity compensation system point upward to the GHe Balloon not visible in the image. The VTI demonstrator segment is retained by tensioned belts and covered with a white protective cloth. Upon release, the retaining 4,2m belt and cover cloth is pushed aside by the VTI 0,8m segments’ hook/loop fastener guided uncoiling motion. The segment is deployed Figure 8. VTI demonstrator deployment sequence 9 International Conference on Environmental Systems

perpendicular to the upper stage, i.e. the GSE trolley, as visible in Figure 8 - (b) until deployment to the full segment length is achieved. At this stage – shown in Figure 8 - (c) – the inflatable beams get fully pressurized which triggers the bending element resulting in a slewing motion of the cryogenic tank covering VTI portion around a predefined axis indicated as dashed line in image (c). With continued slewing motion the VTI segment reaches its final configuration along the cryogenic tank as visible in Figure 8 - (d). The demonstrator had a full length of 5m with a bending axis at 0,8m. The bending angle and final shape of the deployed VTI segment are adjustable by appropriate shaping the inflatable beams which allow the VTI system to closely cover variations in CUS geometries. The VTI demonstrator presented in Figure 8 is largely representative to a flight system. All materials used in insulation and inflatable support structure have been used in spaceflight albeit in unpressurized applications. Demonstrator pneumatic and HDRM subsystems were reduced models made of off-the-shelf components. Still, with full scale demonstrator testing, manufacturability of the VTI segment and the deployment principle of the insulation system are proven.

C. Vibration Testing Vibration testing was performed to verify that the retained deployable insulation is able to withstand typical launch loads without damage. The demonstrator, stowed in rolled form and mounted on a test adapter, was exposed to quasi-static, sine and random vibration. A resonance search was performed before and after each high level test exposure. An image of the VTI demonstrator vibration test setup, showing facility, test adapter, test model and the retaining system is presented in Figure 9. The following test parameters were applied in X, Y, Z axis on the stowed VTI demonstrator element:  Sine Vibration: sweep run with a constant acceleration of 1.2g from 5Hz to 100Hz and a sweep rate of 2 octaves per minute  Quasi Static Test: vibration with a constant acceleration of 4.5g at a constant frequency of 20 Hz for a duration of 30 seconds  Random Vibration: at a level of 20g RMS from 20Hz to 2000Hz for a duration of 4 min/axis Generally, the stowed inflatable insulation showed negligible response against the applied load levels. The rolled demonstrator reveals characteristics of a dampening system and with the given retention design the package was well attached to the support structure. Visual inspections showed no degradation of the packed VTI-MLI structure after vibration testing in all three axes. Resonance search comparisons before and after high level test in all three axes also showed no mechanical degradation in the stiffness of the vibration test setup.

VTI demonstrator Retaining belts stowed

Vibration Test Vibration Test Adapter Facility

Figure 9. Vibration test setup

10 International Conference on Environmental Systems

After vibration testing, another deployment of the VTI segment was performed with the setup described in Figure 8. Unfolding worked flawlessly and the demonstrator did not show any critical degradation due to vibration. Detailed visual inspection on the deployed system only revealed small scale, local abrasions of VDA coating on the MLI outer layer caused by the relative movement during the vibrations. Due to the small size these defects have insignificant influence on VTI performance. In total the current demonstrator model has since been inflated and deployed more than 20 times. Compared to a flight model, which will undergo one test campaign on ground and one deployment in space, the demonstrator was exposed to significantly higher deployment cycles without showing any signs of relevant degradation proofing the robustness of the developed system. Based on successful completion of the above test campaign TRL4 was achieved for the VTI system.

V. Summary and Outlook Within the ESA FLPP program RUAG Space has successfully developed a truly versatile thermal insulation system for cryogenic launcher upper-stages. The pneumatically deployable VTI system is self-sustained and mounted via interchangeable segments on the payload interface adapter. At this location the system is protected from environments on the launch pad and aerothermal loads during lift-off. On the interface adapter the VTI is independent from the upper stage, which in turn does not need to be modified for the insulation system. In-orbit VTI segments deploy upon command in a sequence that concludes in an insulating skirt closed over the cryogenic stage. The entire deployment process is passive and only controlled by built-in retainers made of hook/loop fasteners. Apart from the HDRM assembly no additional mechanism is needed for deployment. This approach reduces system complexity and leads to a mass less than 80kg for a 75m2 covering VTI system. The deployment process has been successfully verified with a full scale, 1.5 meter by 5 meter, model. It was demonstrated that the process is reliable and reproducible in real size and that stowing in rolled form is compatible with launch loads. Thermal analyses have shown that on a generic ø4.4m upper stage the VTI system reduces the heat input into the liquid hydrogen tank by more than 90% resulting in a strongly reduced fuel boil-off rate. VTI may further be developed to TRL6, funding permitting, by performing major system verifications steps such as a 0g flight- and a thermal-vacuum test campaign. As the insulating VTI segments are adaptable in size, the system can be designed to almost any cryogenic upper stage configuration. VTI is thus a viable upgrade kit for existing launchers and future launchers do not need to foresee the system in their initial design. Application of the system may be envisaged for cryogenic upper stages in- development like for the future European launchers Ariane 6 or Vega-E, or the American Vulcan rocket.

Acknowledgments The authors wish to thank the European Space Agency and its Future Launchers Preparatory Program for funding this development within FLPP Periode2/Step 2/ CUST 1.3.

References 1Dye, S. A. et. al., “Design, fabrication and test of Load Bearing multilayer insulation to support a broad area cooled shield,” Cryogenics, Vol. 64, 2014, pp. 135-140. 2Miyakita, T., et. al., “Development of a new multi-layer insulation blanket with non-interlayer-contact spacer for space cryogenic mission,” Cryogenics, Vol. 64, 2014, pp. 112-120. 3McLean, C., et. al.,”Simple, Robust Cryogenic Propellant Depot for Near Term Applications“ 2011 IEEE Aerospace Conference, IEEE2011-1044, Big Sky, MT, 2011, pp. 1-24 4Dew, M., et. al.,”Design and Development of an in-space deployable suns shield for the Atlas Centaur“ AIAA SPACE 2008 Conference & Exhibition, AIAA2008-7764, San Diego, CA, 2008, pp. 1-11 5Freeland, R. E., Bilyeu, G. D., Veal, G. R. and Mikul, M. M., “Inflatable deployable space structures Technology summary”, International Aeronautical Federation, IAF-98-1.5.01, 1998 6Block, J., Straubel, M., and Wiedemann, M., „Ultralight deployable booms for solar sail and other large gossamer structures in space“, Acta Astronautica, Vol. 68, 2011, 984-992 7Moser, M. and Hoidn, W., RUAG Space GmbH, Vienna, Austria, European Patent Application for a “Pneumatically deployable thermal insulation for cryogenic tanks” No. 154510017.7-1754, filed 09 April 2015 8ESA/ADMIN/IPOL(2014) 2, “Space Debris Mitigation Policy for Agency Projects” Paris, 28 March 2014 9ISO 16290:2013 “Space systems -- Definition of the Technology Readiness Levels (TRLs) and their criteria of assessment”, International Organization for Standardization, SO/TC 20/SC 14, Geneva, 2013

11 International Conference on Environmental Systems