1 MECHANISMS

1.1 Requirements and Design Drivers The general requirements assumed for this EU-LISA study is to critically review and re-use where appropriate the design and technology developments already performed in the frame of LISA Pathfinder and LISA mission formulation studies. This will guarantee a TRL level as high as already achieved within those missions, without need to perform additional unnecessary TDAs. Where opportunity for design simplification or mass savings are identified, those would be explored.

Of course maximum level of commonality between the “Mother” spacecraft “MSC” and the Daughter spacecrafts “DSC” shall be pursued.

1.2 Assumptions and Trade-Offs Mechanisms units are identified in the payload, in the spacecraft and at separation between spacecrafts and launcher adaptor.

Specific assumptions and dedicated trade-offs are reported in respective mechanism design sections.

1.3 Baseline Design

1.3.1 Payload

Concerning the payload, a full re-use of mechanisms design / technology developments already performed within Lisa PF and LISA MF is assumed. Those will include Point Ahead Angle Mechanisms (PAAM), Fibre Switching Unit, telescope tracking mechanisms and hold down-release systems.

Figure 1-1: Fibre Switch Unit and Point Ahead Angle Mechanisms as developed

within LISA MF by TNO (NL) and RUAG (CH)

1.3.2 High Gain Antenna Deployment (ADM) and Pointing Mechanisms (APM)

The mission requires to steer a High Gain Antenna around two orthogonal axes.

1.3.2.1 Main requirements • X-band frequency operation

• Two axis steering

• Reflector dish based of 750 mm diameter.

1.3.2.2 Main assumption

Due to reasons explained in section 1.5, the assumption is that the antenna might be built in such a configuration as to have the Centre of Gravity (CoG) close to the APM rotational axes, remaining within a sphere about 50 mm radius during antenna steering. This is assumed to be achievable with on-set type of antennas, plus balancing mass as necessary.

Due to the need to guarantee the necessary clearance the antenna will be mounted on a fixed short boom (likely close to 0.6 m).

1.3.2.3 Baseline design

The HGA Antenna Pointing Mechanism (APM) will mainly be built on the heritage of Rosetta, Bepi Colombo high gain antenna.

The APM will include:

• Two axes (azimuth / elevation) motorized hinges

• Two axes radio frequency rotary joint

• Position sensor

• Hold-down / release mechanisms

• Short boom (0.6 m approximately)

• Thermal hardware (and heather as needed)

• APM drive electronics.

1.3.2.4 Potential problem areas

1) Antenna pointing induced back-action:

Antenna pointing motion generates reactions on MSC / DSC, which might disrupt or degrade laser links; e.g. to 6 deg antenna rotation correspond about 6500 nrad of S/C counter-rotation, which is unacceptable, since to maintain laser link a max of 10 to 100 nrads are required.

Pointing motion approach is based on similar concept as proposed by ASD (RD[1]) , i.e. pointing achieved by series of small angular antenna motion, followed by stabilization phase performed by DFACS.

Similar mass/inertia for S/C – HGA as per previous LISA MF study assumed, which is not strictly true, also because MSC is “heavier” than DSC, but reasonable as a starting point and conservative, since both MSC and DSC show higher masses than in LISA MF.

The following table is extracted from RD[1]:

Table 1-1: Summary of re-pointing Options

It can be seen that stabilization time is given to be proportional to back-rotation to be compensated, thus total travel time is independent from selected minimum step size.

Reference design for HGA pointing mechanism, based on Rosetta / Bepi Colombo HGA APMs heritage (with due design modifications), can achieve resolution down to approx 0.05 deg (approx 50 nrads), thus approx 10.0 sec would be required by DFACS to stabilize after each step.

It has to be noted that under above circumstances the max antenna pointing or re-pointing speed would be 6/22 deg/min (<3 deg /min, ~0.05 deg/sec).

2) Self gravity and antenna CoG motion

System level self-gravity requirement, as applied at antenna level demand that antenna CoG should remain within approx 50.0 mm off-set from rotation axis (RD[1]).

Previous LISA MF configuration assumes only one antenna pointing axis (azimuth), for which rotational axis alignment through mobile parts CoG does not appear a major problem, whilst for this study two axes are necessary.

To maintain the CoG at a reduced off-set from “two” rotational axes requires a detailed study of possible antenna / APM configuration, or ad-hoc to be developed APMs with virtual rotational axes going across antenna mobile parts CoG shall be studied (see Option 2A and 2B).

For the purposes of the present study, the possibility of a compact antenna with CoG motion compatible wrt 50.0 mm radius sphere requirement is assumed.

1.3.2.5 Overall mass estimate:

Overall mass of APM, including all elements as described above under baseline design and under above assumptions can be estimated to be : 10.0 kg (10% design margin to be added).

Figure 1-2: Rosetta Two Axes Antenna Pointing Mechanism (RUAG – CH)

Peak power estimate: 14 W , average power depending on APM duty cycle.

Operational temperature rage, approx: -20C +60C

1.3.3 Telescope Cover & Ejection Mechanisms

Due to cleanliness reasons, a temporary cover for the telescope assembly has to be foreseen for the launch phases. Since the payload and telescope assembly are assumed unchanged w.r.t. LISA Mission Formulation study (at least for the purpose of the telescope cover), same approach and baseline design as identified by ASTRIUM are

proposed (after having checked feasibility and consistency) assumption: derived from LISA MF study RD[2].

The selected baseline is based on a light cover, released by means of a spring actuated mechanism, and a “free hinge” concept, as depicted in here below figure, derived from RD[2].

Figure 1-3: ASTRIUM Telescope Cover Release Concept RD[2]

Overall mass was estimated to be 4.5 kg for each cover plus associated mechanisms.

Two covers would be needed for the MSC, only one for each of the two DSC.

1.3.4 Spacecraft – Launcher Separation Mechanisms

Whilst the MSC will be launched in a single launch, adopting standard adaptor and release mechanism, the two DSCs will make use of customised adaptor and customised release system.

Under current configuration and structural baseline, four hard points will be available for each DSC to implement hold-down possibilities during launch phases. No additional structural interconnection and associated launch-lock / release system, between the two DSCs is foreseen in present configuration / structure baseline.

The overall configuration is very similar to the one studied by ASTRIUM within the “Alternative Launch Configuration “ (alternative to the “Tuna Can” concept). The three spacecraft were accommodated according to the following figure, extracted from RD[2]:

Figure 1-4: “Alternative LISA Launch Configuration – from ASTRIUM study RD[2]

Being located in four corners of the DSC structure, the four hold down points will be able to carry the launch loads. A typical cup/cone system was preliminary envisaged, which is considered appropriate for the current configuration as well (see figure here below from RD[2]).

Figure 1-5 Typical Cup/Cone attachment geometry – from ASTRIUM RD2

Pyrotechnics devices can be assumed as commercially available to perform the release function.

A simple spring driven release concept can be used to guarantee the required trajectory, needed to avoid collision risk during separation (see figure here below from RD[2]).

Figure 1-6: Collision avoidance concept during separation – from ASTRIUM RD[2]

For the present study, an additional 15 degrees titling of the Spacecreft to launcher interface plane is foreseen for accommodation aspects. This is considered fully compatible with above approach for launch load transfer, since the presented concept allows for simultaneous application of longitudinal and lateral loads (as this would be the case for tilted interface plane), and the cup/cone geometry will be suitably oriented to implement the required release trajectory.

Under above assumptions, and for the purposes of a preliminary estimate, 4.0 kg “release mechanism” mass (for each interface point) can be considered applicable (with a 20% design margin to be added).

1.4 List of Equipment 1) Payload associated mechanisms, included in Payload / Instrument budgets (not duplicated here)

2) HGA APM (one per S/C)

3) Telescope cover and ejection mechanism (2 per MSC, 1 for each DSC)

4) S/C Launcher separation mechanisms (only for Daughter S/C)

According to baseline, the following “Equipment Summary Table” has been extracted from the mechanisms work-book.

Table 1-2: Equipment summary from “Mechanisms” work-book

1.5 Options With reference to the HGA antenna pointing mechanisms baseline, the following options can be envisaged:

Option 1

In case the antenna flip over motion would not be possible (due to mission operation constraints), then a continuous 360 degrees azimuth rotation system would be implemented, which would require the presence of slip rings for power / signal transmission to the elevation stage (replacing cable twist capsule).

Impact for this option would be a minor development related to the slip-ring power / signal transmission system to be derived from solar arrays domain.

Option 2A

In case the antenna CoG could not be maintained in a 50 mm radius sphere during steering, than the two rotational axes should be made crossing (or close enough) to the antenna CoG. This would not be a problem for the azimuth axis, for the elevation mechanism a conventional stage would amplify CoG motion during steering. A different type of kinematic, based on converging flexible blades, would be required, realising a “virtual rotational axis” elevation stage. The flexible blades would converge toward the antenna CoG, thus achieving the proper location of the rotational axis. The word “virtual” is used since the rotational axis would be outside the mechanical envelope of the mechanism itself.

Due to the location of the elevation rotational axis, a flexible section of RF W/G shall be used. This is considered feasible, provided the angular range of the elevation axis would remain limited (14 deg range is considered feasible), to avoid overstressing the flexible blades.

Since the initial antenna deployment function (after HRM release) could not be anymore performed by the elevation stage (the angular range would be too large), a

deployment mechanism would be required. It is considered that a simple spring actuated deployment mechanism at the root of the antenna short boom would be adequate. Of course one additional RF RJ (one axis only) and cable wrap, would be needed at boom deployment mechanism hinge level.

This Antenna Deployment Mechanism (ADM) would include:

• Spring based deployment hinge

• Radio frequency rotary joint (one axis)

• Cable wrap to route power / signal through ADM toward APM

• Associated thermal hardware (passive type).

Figure 1-7: Spring Actuated Appendage Deployment Mechanisms (Astrium)

Impact for implementation of this option would be:

a) Development of an elevation stage based on “virtual rotational axis’, that is based on flexible blades converging toward the antenna COG. This development is estimated to require a ROM of 500-800 kEuro.

b) Implementation of an antenna deployment mechanism, including one additional RF rotary joint, for a total added mass of about 6.0 kg.

c) A flexible W/G section at elevation axis level replacing one axis of the RF rotary joint from baseline design configuration.

Option 2B

Under the same assumptions as Option 2A, another interesting concept could be to implement the deployment mechanisms at antenna dish level. It cannot be excluded, indeed that an antenna configuration where only the reflector dish can be stowed and further deployed, instead of involving the entire RF chain. In this case a simple spring based mechanism (simpler and lighter than the one from option 2A) could be placed between the reflector dish and the elevation stage of the mechanism.

Impact for implementation of this option would be one additional (small) reflector deployment (and locking) mechanisms, for a mass estimate of 2.0 kg.

Option 3

Again in case the CoG containment within a 50 mm radius sphere could not be guaranteed, still the mechanisms baseline design could be used, provided some balancing masses could be suitably placed to bring indeed the overall antenna CoG in the desired location.

It is obvious that this option would imply a mass penalty (assumed to be less than 6 kg at present), and a detailed study to position the balancing masses within permitted envelope to avoid collision risks with S/C body.

Impact at system level would be a mass increase than can be estimated approximately 6.0 kg.

1.6 Technology Requirements The following technologies are required or would be beneficial to this domain:

Included in this table are:

– Technologies to be (further) developed – Technologies available within European non-space sector(s) – Technologies identified as coming from outside ESA member states. No specifc technology development is required for the presented baseline. In case Option 2A would be pursued, the development of a “virtual axis” based

antenna elevation mechanisms should be developed (see details in Option 2A section).

Equipment and Text

Reference

Technology Suppliers and TRL Level

Technology from Non-Space

Sectors

Additional Information

NA

1.7 Acronyms Add your acronyms to the table provided below:

Acronym Definition

APM Meaning of the acronym

ADM Meaning of the acronym

HRM Hold-down release mechanisms

RF Radio Frequency

RJ Rotary Joint

CoG Centre of Gravity

W/G Wave Guide

MSC Mother Spacecraft

DSC Daughter Spacecraft

1.8 References RD[1] « Telecommunication System Technical Note », LISA-ASU-TN-4003

RD[2] « Lisa Alternative Launch Configuration », LISA-ASU-TN-4016