XEUS MSC DOCUMENT - EMITS Invitation To Tender...

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MSC Spacecraft Definition

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f D O C U M E N T

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XEUS MSC DEFINITION DOCUMENT

prepared by/préparé par A. Lyngvi reference/réference SCI-A/2006.013/AL issue/édition 1 revision/révision 0 date of issue/date d’édition 26 April 2006 status/état Final Document type/type de document Technical Note Distribution/distribution

XEUS MSC Definition Document issue 1 revision 0 - 26 April 2006

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T A B L E O F C O N T E N T S

1 INTRODUCTION............................................................................................................. 4

2 TERMINOLOGY.............................................................................................................. 6

3 MISSION ANALYSIS...................................................................................................... 7

4 CONFIGURATION.......................................................................................................... 8

5 STRUCTURE .................................................................................................................13 5.1 Spacecraft Structure .........................................................................................................................13

5.1.1 Service Module Structure.........................................................................................................13 5.1.2 Sun baffle .................................................................................................................................14 5.1.3 Optical Bench...........................................................................................................................14

5.2 Optics covers....................................................................................................................................15 5.3 Mass summary..................................................................................................................................16

6 THERMAL.......................................................................................................................17

7 COMMUNICATION.......................................................................................................18

8 DHS..................................................................................................................................19

9 AOCS...............................................................................................................................20

10 PROPULSION................................................................................................................21

11 POWER...........................................................................................................................23

12 OVERALL MASS ..........................................................................................................24

13 REFERENCES...............................................................................................................25


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1 INTRODUCTION The X-Ray Evolving Universe Spectroscopy (XEUS) mission is a candidate for the next ESA X-ray science mission, presently being investigated by SCI-A as a reference study, in preparation to the forthcoming Cosmic Vision selection process. The baseline design of the XEUS mission provides for a far larger collecting area and significantly improved angular resolution compared to XMM-Newton. To obtain such large collecting area the mission would consist of two spacecraft flying in formation at L2; a detector spacecraft (DSC) and a mirror spacecraft (MSC). These two spacecraft would be separated with a distance of approximately 35 m in the baseline design (figure 1). Following recent changes in the XEUS international cooperation scenario, the overall mission baseline design has been revisited, in order to reduce complexity, development risk and cost. In particular, the baseline design of the Mirror Spacecraft has been modified to assume a fixed optical bench configuration. This document is written in preparation of the XEUS industrial activities. The document describes a possible reference design of the Mirror Spacecraft (MSC). It is emphasized that the design and considerations described throughout this document are preliminary and hence all trade-offs will have to be revisited in order to properly obtain a detailed MSC design. The document can be used as a starting point for industrial and internal design activities and it is assumed that major changes will be identified through these activities hence it is acknowledged that the final design of the MSC will be different from what is described in the text.

+ 20 deg

DSC

Sun

Object to be observed

MSC


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Figure 1 Mirror (MSC) and Detector Spacecraft (DSC) maximum sun aspect angles observing at L2.

Sun

Object to be observed

DSC

MSC

90 degrees sun aspect angle


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2 TERMINOLOGY Sun Baffle (MSC): Cylindrical structure around the Mirror Spacecraft (MSC) protecting the telescope elements from direct Sun illumination as well as supporting the solar arrays, contributing to the thermal control of the telescope and playing an x-ray baffling role. X-ray skirt (MSC): Annular surface located on the external part of the MSC, approximately perpendicular to the optical axis and playing an important role in preventing x-ray straylight to reach the focal plane detectors. It can be fixed (coinciding with part of the optical bench) or deployable, located externally to the MSC envelope. Detector X-ray baffle (DSC): Cylindrical baffle located at the optical aperture of the instrument focal plane (on the Detector Spacecraft), playing an important role in preventing x-ray straylight to reach the focal plane detectors. Mirror module: A mirror module is a stack of about 50 (TBC) ind ividual mirror plates of 250 micron thickness and 300 x 300 mm size (TBC) emulating a double cone approximation to a Wolter I design. It is assumed that during the development phase, the mirror modules will be procured from the mirror manufacturer. Mirror petal: A mirror petal is a group of mirror modules emulating a double cone approximation to a Wolter I design, integrated into a single element structure and equipped with a front and a rear collimator used as X-ray baffles. The petals are assembled on the optical bench. Optical bench: The mirror optical bench is the mechanical structure that supports the mirror petals with the appropriate alignment accuracy in-orbit. Telescope: The XEUS telescope consists of the mirror optical bench equipped with all mirror petals and of all the necessary equipment needed to fulfil the XEUS telescope performance and functional requirements. Hence, the XEUS telescope also includes thermal hardware for thermal control, optical corner cubes and metrology subsystems to be used for formation flying, optical baffles for straylight prevention, alignment devices …etc. The telescope might also contain electron deflectors to prevent low energy electrons from hitting the X-ray detectors. However, the location of such deflectors could also be located on the DSC and thus a detailed trade off on the location will be conducted.


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s 3 MISSION ANALYSIS

The XEUS mission is a formation flying mission located at L2. The operational orbit is assumed to be a Halo orbit as seen in Figure 2. The benefit of this orbit is that almost no manoeuvres are required for insertion at L2 and that almost no manoeuvres are required for orbit maintenance. The XEUS mission baseline is using and Ariane 5 launch vehicle with a direct injection towards an L2 orbit. To minimize operational complexity and to avoid any rendezvous manoeuvres at L2 the detector spacecraft (DSC) and the mirror spacecraft (MSC) will be launched in a stacked configuration and continue travelling towards the science orbit in the stacked configuration. Using this approach the Delta-V requirements on the spaceraft can be minimized and only some minor trimming manouvers and correction of launcher dispersion will be needed for the cruise phase.

Figure 2 Halo orbit around L2 in synodic space: ecliptic projection, Earth at coordinate (0, 0) and Sun along positive x-axis Once in science orbit the two spacecraft would separate and DSC would need to deploy any solar arrays or other appendages before acquiring the formation could start. The formation would be acquired through a set of sensors and manoeuvres using RF and optical metrology. However, in this phase and other formation flying phases the MSC will be considered the master satellite, thus the DSC will be the active spacecraft making sure that it flies in the correct formation. The orbit maintenance for the orbit assumed to be very low and a Delta-V of about 1 m/s a year (without margin) is sufficient. In total the Delta -V requirements and therefore also the propellant requirements for the DSC are very low.


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4 CONFIGURATION Previous work has been analysing a configuration based on a large deployable mirror structure, conceived to maximise the telescope geometric area for a given launcher fair ing volume. Such a design imposes considerable challenges in terms of deployment mechanisms, overall MSC configuration and related qualification and testing. In addition, it proves as beneficial to maximise the mirror effective area only when it is possible to populate the entire available volume, i.e. in presence of adequate launcher performance, presently exceeding the expected A5-ECA performance (final injection at L2). Based on these considerations and on the recent programme evolution, it has been decided to explore an alternative MSC configuration based on a fixed optical bench, with a circular geometry, described in this document. The main design drivers to be taken into account are:

- maximize telescope area (crucial to meet the science requirements) - meet the strict alignment requirements imposed by the optical design - provide a controlled thermal environment to telescope optics - accommodate all subsystems required by autonomous MSC operations - compatibility with resource envelopes (mass, fairing size) - compatibility with launch configuration (MSC+DSC stack, launch environment, including

protection from contamination of the optical surfaces) - compatibility with formation flying requirements

Using a fixed optical bench effectively constraints the maximum mirror size to the fairing diameter. However, as the X-ray mirror works as a focusing lens, a considerable part of the available area to the mirror would need to be used by the spacecraft. Therefore the actual mirror area will be less than the area of the fairing cross section. A potential configuration of the XEUS stack in the launcher is shown in Figure 3. This configuration is driven by maximizing the effective area of the mirror. This maximization of the mirror area led to the configuration of MSC, where the MSC spacecraft bus, which contains all the relevant subsystems would be located in the centre of the mirror. Surrounding the mirror is a Sun baffle that prevents Sun light from hitting the mirror module, the optical bench or the inner service module. The size of this baffle is driven by the clear need of having a portion of the sky accessible for measurements at all times during the mission. The pointing direction approximately perpendicular to the Sun direction that would be accessible with the configuration as shown in Figure 3 is constrained to a total range of 20 degrees, with 20 deg outwards the Sun direction and 0 deg in the Sun direction (Figure 1). The definition of such a pointing range is related to the preference for a fixed baffle and to the need of avoiding any interference with the launcher. As a result, the allowed Sun aspect angle in one direction is much la rger than in the other, due to the short distance between the separation plane in the launcher and the mirror. Currently it is assumed that the spacecraft will not be able to off-point


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s in one direction. The actual off-pointing angles in both directions wil l be consolidated by dedicated thermal and straylight analysis and subject to a specific consolidation process during the project evolution.

Figure 3 To the left, configuration of the MSC in the launcher with DSC on top. On the right, figure showing the different baffle measurements for a baffle compatible with a +15/-5 degree off-pointing. As it is assumed that the mirror is located about 0.5 m above the spacecraft lower panel, the max baffle length on this side would be about 0.61 m. In the other direction, the baffle size will be less constrained and the current assumption is that an off-pointing angle of 20 degrees is possible (again this will have to be consolidated based on thermal and structural analysis). The maximum baffle length in this direction would then be about 3.68 m. Maximisation of the telescope area is crucial in order to meet the mission science requirements. Given the proposed configuration, maximisation of the area can only be achieved by optimising the design as to minimise the central volume occupied by the S/C subsystems (impacting on launcher


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s adapter selection) and the radial length required to support the Sun baffle and any related functional units, such as thermal control and solar array (impacting on structural design and service accommodation).

Figure 4 MSC dimensions in the vertical plane in the launch configuration in the sun ant-sun direction Given the rather limited margin available to maximise the radial length available to optics in the direction of the Sun baffle, adapter selection is a crucial point. To have a simple structure it would be beneficial to use an off-the-shelf adapter with a simple connection to the main structure. In the configuration presented in Figure 4 the 1194H adapter is used. The structural analysis performed indicates that this adapter could sufficiently sustain the loads from the two spacecraft during launch. Thus the baseline is a ~1.2 m diameter main spacecraft structure. This allows an inner diameter of the mirrors of close to 1.3 m. To be compatible with the small 1194H adapter and because of the total mass of the stack is rather high (~6500 kg) the stack needs to keep the centre of gravity rather low (Figure 1). In the current


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s configuration this is obtained by having the MSC located on the bottom of the stack with the mirrors mounted as close to the separation plane as possible.

Figure 5 1194H load capability The outer diameter of the baffle is driven by the faring diameter of 4.57 m and it is currently assumed that the outer diameter extends out to a diameter of 4.34 m thus allowing some room for accommodation of external components such as thrusters, solar arrays etc. In the current configuration an x-ray skirt with a width of 11.5 cm is mounted on the outer face of the sun baffle, which results in an 16 cm skirt when accounting for the area just inside the sun baffle that will also function as an X-ray skirt. In addition the extension of the sun baffle will also function as an X-ray skirt. The sun baffle extends in the direction of the optical axis, but will be offset by a small angle and thereby it will limit the solid angle of the detector to space. For a 1.25 m baffle, the X-ray skirt length contribution would be about 7.5 cm, thus leading to an effective skirt size of 23.5 cm. This is shown in Figure 6. A trade-off between baffle on the DSC and skirt on the MSC can be performed to optimise the actual skirt size where a larger skirt would lead to a smaller baffle on the MSC. It is currently envisioned that a skirt size of 23.5 cm is a minimum and an increase in skirt size would be beneficial. This could for instance be done with a deployable skirt. The possibility of such a skirt would have to be investigated in further detail. A major difficulty for the definition of the X-ray skirt is the need for having a rather thick metallic layer to stop the incoming x-ray photons. The thickness of such a skirt should be thick enough to provide 1 attenuation lengt h at 17 keV, which corresponds to a thickness of 722 µm of Al or 4.5 microns of Au. Obviously for the sun baffle the incidence angle of any photons would be very large thus the coating of this part would not need to be as thick.


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Figure 6 Illustration showing the effective skirt length


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5 STRUCTURE

5.1 Spacecraft Structure

The structure in the accommodation above consists of three main parts, the optical bench, the sun baffle and the service module structure. A drawing showing the dimensions of the structure in one plane is shown in Figure 4 while the structure supporting the mirror petals (the optical bench) and its dimensions is shown in Figure 5. The service module is in this configuration defined as the structure of inner cylinder, which will contain the spacecraft elements such as data handling, propulsion system etc.

Figure 7 Top view of the spacecraft structure and its dimensions

5.1.1 SERVICE MODULE STRUCTURE In the current baseline the spacecraft subsystems would be located in the inner cylinder that would have a diameter of about 1.2 m. This allows a direct load path from the adapter through the main structure up to the DSC interface. The current height of this cylinder would be about 2.5 m. To ease the integration it was decided to have 2 equipment shelves mounted in the horizontal direction. All spacecraft components would be mounted on these shelves. The propulsion tank is mounted close to the launcher adapter interface, while the rest of the spacecraft equipment is mounted towards the DSC interface side. The current assumption is that these equipment shelves


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s would be about 1 m wide. If we currently assume that the surface density of the main spacecraft structure would be about 8 kg/m2 and that 20% mass is required for additional secondary structure the total mass of the main spacecraft structure would be about 166 kg including 20% margin.

5.1.2 SUN BAFFLE The sun baffle would be surrounding the optical bench of the mirrors and would be shaped so that no direct sun illumination would impinge on the mirrors. In order to avoid direct sun reflection from its structure to the optics at sun aspect angles from 0-20 degrees, the baffle needs to be extended as high 3.7 m on the sun side and 2.1 m on the anti sun side. The sun baffle will need to be optimized also from a thermal point of view as it is of great benefit to have the mirror at a temperature close to room temperature. This would most likely imply that the sun baffle would need to be of a conductive material potentially even embedded with heat pipes. The current baseline for the sun baffle is therefore based on Al. structure. As the sun baffle itself is not a load bearing structure it is assumed that a surface density of about 4 kg/m2 is sufficient. This would then imply a total structural mass of the baffle of about 190 kg including 20% margin in the current configuration. An X-ray skirt would also be mounted on the sun baffle. This would be in the form of an annular disk with a radius of 16 cm from the outer edge of the sun baffle. To properly attenuate the x-ray it is assumed that in worst case a 1 mm thick aluminium sheet would be needed. This would add another 6 kg.

5.1.3 OPTICAL BENCH The mirror structure is driven by the optical requirements of the mirror and by the need to support its optical elements (called ‘petals’). The mirror structure will function as an optical bench and as the mirror has very strict misalignment requirements this structure will have to be carefully designed both with respect to thermal expansion and testing on ground. The thermal expansion, the stiffness and the compatibility with the actual petal structural design will drive the material selection of this mirror optical bench. The mirror elements consist of individual petals (Figure 8), each containing a number of sub-units (referred to as ‘tandems’), to cover a given fraction of the total area. The petals are presently under development under a separate ESA contract. Presently it is assumed that the external structure of the petal will be constructed using ceramic materials such as CeSiC.

Figure 8 Preliminary drawing of a single petal of the XEUS mirror.


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s It is currently assumed that the petal is large enough to cover the entire area from the inner to the outer radius (i.e. from ~0.6 m to ~2 m). This assumption needs to be verified in the context of the HPO development study as it is expected to impact on technical as well as programmatic issues. In case of limitations on the maximum radial length of each petal, it will be necessary to adopt two concentric rings of petals, with associated penalties in terms of obscuration effects and effective area. Any impacts on the mirror optical bench also need to be investigated. Petals are supported by radial beams (spokes) departing from the central MSC structure, which are also used to stiffen the optical bench and overall system structure (see Figure 7). More detailed interfaces can be found in RD-1. It is currently assumed that 16 such spokes are used (to be confirmed by dedicated mechanical analysis). Each spoke is assumed to have a width of about 8 cm. The width of these spokes will affect the effective area of the optics, therefore having these spokes as thin as possible is an advantage. The spokes height however will not affect the effective area significantly, in fact if these spokes are high enough they may be used to limit the petal’s view factor to space and therefore be beneficial from a thermal point of view. If each spoke has a wid th of 8 cm then the petal outer arc will be about 77 cm and the inner arc would be about 19 cm. Both at the inner radius and the outer radius a support ring of equal dimensions to the radial beams are required. With the assumption that the height of all the elements is approximately 0.5 m the total mass would be about 413 kg including secondary structure and margins.

5.2 Optics covers

The optics of the XEUS MSC will have strict contamination requirements in particular particulates contamination. The faring provides a class 100 000 facility, but during flight contamination will arise from outgassing and from thruster firing. This means that the MSC will have to use covers to limit this contamination. Two covers are needed for the MSC, one on the back side and one on the front side. These covers need to be deployed (or jettisoned) probably after separation from the DSC and after an initial outgassing period. Several alternatives exist on how to implement this and some two of these are described below. When choosing an appropriate cover design minimizing the complexity is of paramount importance. 2 cover solution: A 2 cover solution where each cover covers each of the two sides might be a

simple solution to this issue. The covers could then either be jettisoned (although this might not be possible for debris avoidance reasons) or it could be hinged, with one hinge on the anti-Sun side of the spacecraft and thus be located in the shade of the spacecraft after opening. However using this approach also has some major drawbacks. Firstly, this approach is likely to lead to a cover where also major parts of the spacecraft would be located under the cover. The impact of this would then need to be carefully assessed as it might put restrictions on the use of MLI on the outside of the spacecraft structure. Both the thermal impact and the cleanliness impact of this approach need to be investigated in further detail. Secondly, by having such a large cover it is likely that in order to withstand the necessary loads a stiff structure, such as Aluminium honeycomb would need to be used. This would then cause a higher mass of the covers.


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s Covers opening inside the drum: A solution where the covers would hinge open inside the drum

would also be possible. This might be done by having covers attached with hinges on the spokes. Using this approach there will be a need for having some clearance on the inner arc of the cover towards the central cylinder as when moving into open position the cover would need to be straight at the inner part instead of arcing. If there are 16 covers, one for each spoke, the loss in covered area would only be very small. Another drawback with this solution is that it will require a certain baffle length so that the covers are not Sun illuminated when off pointing. In the current configuration this implies that only the side with the large Sun baffle would be able to use this type of covers.

There are many other cover solutions, such as covers on a rail, thin permanent covers that would be transparent to x-ray etc. However, these solutions are presently considered as more complex or in general less feasible. When estimating the mass for the covers, the solution where the covers are opening inside the drum is selected as this will likely lead to a higher mass as the structure needs to be stiff and hold down and release mechanisms (HRM) for 32 covers are needed. As there are quite some uncertainty on how the covers would look like, the mass of them has been calculated using a structure that would for sure be strong enough. Using 3.5 kg/m2 implies a typical spacecraft structure made of Aluminium, if this is used to cover the two sides then a mass of about 124 kg including margin is required. In addition HRM and deployment hinges are required for the covers, assuming 2 kg per cover the total mass of the cover mechanisms would be 64 kg (incl margin).

5.3 Mass summary

Based on the discussion above the following mass budget is derived for the MSC structure. Mass (kg) Margin Mass incl. margin (kg) Total SVM structure 138 20% 166 PLM Sun baffle/drum 158 20% 189 Optical bench structure 344 20% 413 X-ray skirt 6 20% 7 Covers 104 20% 124 Cover mechanisms 53 20% 64 Total PLM 664 797

Table 1 Structural and mechanisms mass for MSC


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6 THERMAL The biggest issue of the thermal control is the temperature gradients of the mirror modules. The temperature gradient requirements of the mirror are very strict due to alignment issues (RD-1) both in the radial direction and in the direction of the optical axis. To minimize the thermal gradients it is currently envisioned to transport heat from the body mounted solar arrays on the sun side to the antisun side. In the current configuration the sun baffle would effectively limit the view factor to space, thereby increasing the temperature which should preferably be as close to room temperature as possible. Another way of decrease the gradients and to increase the mirror operating temperature is to use the heat dissipated in the spacecraft main structure, which is located in the middle of the mirror to transport the heat to a suitable location (likely to be on the antisun side). Additionally, the x-ray baffling done on petal level should also have a very important thermal function and coating and design of this baffle design will likely be driven by these considerations. The sun baffle would likely need a quite even temperature distribution around the drum. This is currently envisioned through use of different external coatings on the antisun and the sun side; a low emissivity coating on the antisun side and a high emissivity coating on the sun facing side. Additionally it is currently assumed that heat pipes are embedded in the structure. The number of heat pipes will definitely undergo optimization, but currently about 40 m of heat pipes are estimated, leading to a mass of about 17 kg including margin. The thermal control of the optics will be investigated in further detail when the design of the mirror has sufficiently matured. For now the thermal control of the optics has an allocated mass of about 50 kg. This could potentially increase substantially if structures are required for thermal reasons in order to limit view factor to space. The table below shows the thermal hardware as estimated in the baseline design. Mass (kg) margin Mass with margin (kg) PLM heat pipes 14 20% 17 MLI 13 20% 15 paint 8 20% 9 Thermal control optics 42 20% 50 Total PLM 76 20% 91 SVM MLI 6 20% 7 Misc 3 20% 4 Total SVM 9 20% 11

Table 2 Thermal Control Mass budget


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7 COMMUNICATION The MSC communication system will in the baseline design do the communication with Earth during the cruise. Thus the telecommunication system would need to be designed in order to transmit the housekeeping data for both MSC and DSC. It is currently not envisioned that the MSC will be able to transmit any of the science data, as this will be done by the DSC, thus the total amount of data that needs to be transmitted assumed to be not more than about 5 kbps. The current assumption is that the New Norcia ground station will be available for the XEUS mission. Using the 35 m dish it will be no problem to achieve the required data rate with a low gain antenna operating in X-band. Another function of the telecommunication system is the metrology system. This system is based on the work done for the Darwin study in which and S-band system is used to determine the relative position of the two spacecraft in respect to each other. The system can also be used as a communication link between the two spacecraft (~ 9kbps) and therefore it will be able to provide redundancy of the downlink to Earth. The RF metrology system will be the same for both the DSC and the MSC and it will be able to provide a precision similar to what is shown in Table 3.

Table 3 RF metrology precision Based on the above discussion the typical mass of the telecommunication system will be as described in Table 4. Mass (kg) Number of units margin Mass incl. Margin (kg) X-band LGA 0.1 3 20% 0.36 X-band transponder 3.45 2 20% 8.28 X-band SSPA 1.3 2 20% 3.12 X-band RFDU 1.5 1 20% 1.8 S-band transponder 3 2 20% 7.2 S-band antenna 0.2 4 20% 0.96 S-band RFDU 0.3 1 20% 0.36 S-band patch antenna 0.1 2 20% 0.24 Total 22.32

Table 4 Mass budget for the communication system


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s 8 DHS

The XEUS DHS shall contain the interface to the AOCS, handle the telecommands, acquire telemetry, communicate with the DSC, perform power and thermal control, provide on-board timing and storage of on-board house keeping data and failure detection and recovery functions. A possible des ign could be based on the ongoing development of the Highly Integrated Control and Data System (HICDS) and using this approach the DHS could be as indicated in Table 5.

Figure 9 DHS baseline design The mass memory required for the XEUS MSC is very limited as it is only required to store housekeeping data. The use of the HICDS is only one alternative of possible CDMU for XEUS MSC. In the current configuration the actual process ing needs are not excessively high and thus the selection of the DHS should be based on technology maturity and cost. However, to get an estimation of the resources required the DHS the HICDS is used, in which the mass budget is provided in Table 5. mass (kg) margin mass incl. Margin (kg) CDMU + Mass memory 6 20% 7 I/F 2 20% 2 Total 7 9

Table 5 Mass budget for the DHS

CDMU

PCDU TCS RCS RW STR CSS Gyro

Command & Control Bus(1553 or CAN)

X-bandtrsp

S-bandtrsp


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9 AOCS The Attitude and Orbit Control System (AOCS) is will provide the required spacecraft attitude of the MSC in science orbit and of the stack during cruise. The actual pointing requirement for the MSC is rather relaxed with an absolute pointing requirement (APE) of approximately 60 arc seconds. When flying in formation with the DSC, MSC will be the master in which the DSC will follow the movements of the MSC. Thus the formation flying part of the AOCS for MSC is likely to not be very complicated. A traditional AOCS would therefore be sufficient for the MSC. In the current design the mirrors are located towards the separation plane in order to keep the CoG as low as possible during launch. The drawback of this approach is the fact that the centre of pressure and the centre of mass would likely have a large offset. In the current calculation the offset is assumed to be approximately 1 m. For a spacecraft with a cross sectional area of approximately 20 m2, the SRP would create a torque of approximately 1.5x10 -4 Nm. The longest observation period is approximately 400 ks and hence the reaction wheels would need to have a capacity of about 70 Nms assuming that 1 out of 4 reaction wheels have failed. The sensors on the MSC contain the typical set of star trackers, sun sensors and gyros in which all should be mature off the shelf items. Using typical numbers for these elements we derive the numbers in Table 6 for the AOCS mass budget. mass/unit (kg) number margin Mass incl. Margin (kg) Reaction wheel 15 4 5% 63 Star Trackers 7.5 2 5% 16 Gyro 4.5 1 5% 5 Fine Sun sensor 0.5 2 5% 1 Coarse Sun sensor 0.2 3 5% 1 Total 85

Table 6 Mass budget for the AOCS system


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10 PROPULSION The propulsion system is designed for a lifetime of 5 years nominal and 5 years extended. This implies that during the 5 nominal years the full set of margins will apply, while for the extended lifetime no margins will be added. The propulsion system is currently based on a monopropellant hydrazine system. This is selected over cold gas for its much higher efficiency. The major drawback of such a propulsion system is the possibility for contamination coming from the thruster. To avoid contamination the thrusters would need to be mounted in such a way that the thruster plumes would not directly impinge on any mirror surface. Additionally, the thrusting in the direction of the DSC should be avoided. This clearly puts constraints on the thruster accommodation on the MSC. In the current design the thrusters are located on the outside of the sun baffle and thus they have no direct view to the mirrors. Nevertheless, as the mirrors are very sensitive to contamination it is clear that an analysis of the contamination coming from the propulsion system is required. The baseline design uses 10 N hydrazine thrusters to provide the required off- loading. These thrusters would typically be located at least 2 m from the CoM. Over a time of 5 years the thrusters would then need to provide a total impulse of about 11 kNs in order to counter act the solar radiation pressure. With the appropriate margins over the entire extended lifetime of 10 years, the total impulse requirement would be 34488 Ns. In addition the propulsion system would need to provide the required delta - v for orbit maintenance, launcher dispersion etc. In total the de lta-v requirement is about 50 m/s. Impulse Total impulse (nominal) 11495.92 100% 22992 Total impulse (extended) 11495.92 0% 11496 Total 34488 Delta-V Delta-V launcher dispersion 30 5% 32 Delta-V cruise 3 5% 3 Delta-V orbit maintenance (nominal) 5 100% 10 Delta-V-orbit maintenance (extended) 5 0% 5 Total 50

The Delta-V of 50 m/s and the total impulse of 34488 Ns results in a propellant requirement of 130 kg. This propellant requirement should be compatible with a hydrazine system in blow down mode, thereby greatly simplifying the propulsion system. This would imply that the tank volume would need to be oversized by 25 % in order to account for the pressurant gas. In addition a 10 % margin on the tank volume is required, which results in a tank volume of minimum 178 litres. The overall mass budget of the monopropellant hydrazine system working in blow down mode is shown in Table 7.


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s Mass/unit( kg) Number Margin Mass incl. Margin (kg) Tanks 16 1 20% 19 thrusters 0.25 16 20% 5 Piping and misc. 6 1 20% 7 total 31

Table 7 Mass budget for propulsion system


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11 POWER To maximize the available mirror area the sun baffle is mounted as close as possible to the faring wall. This implies that there are limited room for accommodation of equipment on the outer surface of this baffle. GaAs will be very complicated to mount on a curved surface, and therefore silicon cells are baselined in the current configuration. Silicon cells will have a much lower efficiency, but are commonly mounted on curved surfaces. As the solar cells are mounted directly on a surface the temperature of the solar arrays will be rather hot, which will decrease the efficiency of the solar arrays. In the current solar array sizing it is assumed that the temperature of the solar arrays will be in excess of 100 °C. However, this would need to be investigated further, in particular as the sun baffle would contain heat pipes that will transport heat to the antisun side. This clearly would decrease the temperature of the solar arrays. Nevertheless, in a worst case scenario the temperature of the solar arrays will be very hot and the efficiency is therefore assumed to be about 5 % EOL. The power demand of MSC is about 450 W which result in a solar array size of about 6.5 m2. With the necessary margin the solar array size would be about 8 m2. An important point to verify is the actual power demand when the spacecraft travels in a stack, as it is likely that the MSC will need to provide sufficient power to the DSC. Depending on the power need of the DSC, this will increase the solar array size of the MSC. The silicon cells are mounted on a separate curved substrate which would be mounted on the external wall of the MSC sun baffle. It is assumed that the substrate would be 10 mm of Aluminium honeycomb with a specific mass of 3 kg/m2, including the solar cells. Additionally, the MSC will require a battery to cope with the peak power demand in science orbit and the Launch and Early Operations Phase (LEOP). The mass breakdown of the power system is shown in Table 8. Mass/unit (kg) Unit margin Mass incl. Margin (kg) solar array 3 8 20% 28.8 battery 8 1 20% 9.6 PCDU 5 1 20% 6 Total 44.4

Table 8 Mass budget for the power system Harness The harness mass should be estimated as 10 % of the dry mass before maturity margin. This in particular is necessary due to the large distance (from the outside of the drum) to some of the components.


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12 OVERALL MASS Based on the descriptions above the following mass budget for the MSC is derived. This shows that a mirror mass of 1500 kg, should be compatible with the current configuration of the MSC. However, as the uncertainty in the design of the MSC is still large, sufficient margins should be applied. Therefore, future design iterations the PLM (i.e. the telescope) shall use a maximum mass envelope of 2700 kg for the entire telescope before system margin (20%). On the basis of similar arguments, a maximum SVM wet mass of 750 kg (including system level margin) should be used. Item Mass wo margin Mass (incl. 20 % design

maturity margin) PLM (telescope) Structure (incl. optical bench) 508 609 Covers 157 188 Thermal 76 91 Mirror 1500 1800 Harness 62 74 Total PLM mass 2302 2762

System Margin 20 % 552 Total PLM mass w. margin 3314 SVM Structure 138 166

Thermal 9 11 Comms 19 22 DHS 7 9

AOCS 81 85 Propulsion 26 31 Power 37 44 Harness 26 32 Total SVM mass 344 400 System Margin 20 % 80

Total SVM dry mass w. margin 480 Propellant 130 130 Total SVM wet mass w. margin 610 Total mass (PLM+SVM) 2775 3925

Table 9 Overall mass budget for the MSC


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13 REFERENCES RD-1 XEUS Telescope Performance Specification and Interface Requirements

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