Status of the ESA Feasibility Study HYPER. Ulrich Johann Astrium GmbH (for the industrial team) CNES, Paris

Transcription

1 Status of the ESA Feasibility Study HYPER Ulrich Johann Astrium GmbH (for the industrial team) CNES, Paris

2 Presentation outline: Scientific Objective of the HYPER Mission The present HYPER Mission Baseline The HYPER Feasibility Study The Lense-Thirring Effect and HYPER HYPER Measurement Principle and Payload Definition HYPER Technical Requirements Classification Orbit Selection Trade-off Payload Configuration Trade-off Atomic Sagnac Unit Conceptual Design PST Conceptual Design HYPER Payload Module Design The HYPER DFAC Simulator Model Drag-free and Attitude Control (2.AOCS) Design (incl. IS, FEEP technology) Summary and conclusions (status, remaining work, mission road map outlook)

3 HYPER scientific mission objectives Mapping the spatial structure of the general relativistic gravito-magnetic effect of the Earth with better than 10% accuracy Independent determination of the fine structure constant to test quantum-electrodynamics theories Investigation of quantum de-coherence to set an upper bound for quantum gravity models Demonstration of the superior performance of cold atom sensors for spacecraft control Technical implications To fully develop this potential, however, it is necessary to: establish the operational environment for atom interferometry by proper spacecraft and payload engineering, namely: drag-free control and supreme inherent stability of payload elements and pointing performance. While the second and third science objective take advantage of the space environment only, the measurement of the Lense-Thirring effect necessitates a low earth orbit. HYPER will also pursue the development of atom interferometry as a high precision sensor for spacecraft control in the future, driving the design towards compactness and robustness.

4 HYPER Mission Baseline; June 2002 Payload Mass: Payload Power: Payload Dimensions: Optical Bench: Atom Preparation Bench: (not a detailed subject of this study) Laser Bench: (not subject of this study) Orbit: Launcher: Launch year: Mission lifetime: Spacecraft launch mass: AOCS: Propulsion: Power: Telemetry: Ground segment: kg 203 W 937 mm diameter x 1300 mm height Optical elements for coherent atom manipulation High-precision star tracker (200 mm Cassegrain Telescope, pointing performance, 10 Hz readout frequency) 2 drag-free proof masses (asymmetric arrangement) 4 atom interferometers based on caesium or rubidium accommodated in 2 magnetically shielded vacuum chambers Optics for atom preparation and detection Lasers for atom interferometry (Raman), preparation and detection of the atoms High-precision µw synthesiser for the hyperfine transitions of caesium or rubidium Dawn-dusk, Sun-synchronous orbit at 700 km altitude, 98.2 inclination, 98.6 min Low-cost Rockot launch vehicle from Plesetsk Cosmodrome To be based on road map results (in terms of technical readiness, 2009) 2 years (nominal) kg (Launcher capability 870 kg to 700 km) Primary AOCS + secondary AOCS using Payload Module sensors for error generation during Science Mode (drag-free control and fine pointing) 16 x 500 µn FEEP thrusters + 8 x 40 mn cold-gas thrusters 499 W (EOL) Fixed 3.3 m 2 GaAs solar array + 6 Ah Li-ion battery S-band, 500 kbps during 7-minute passes, total of 190 Mbit/day 15 m-antenna at the ESA Kiruna station, Mission Control ESOC

5 The HYPER Feasibility Study Study overview ESA/ESTEC Invitation to tender February 2002 Proposal March 2002 Study Kick off June 6th, 2002 Intended study duration: 6 months Final presentation planned for February, 2003 Study team: Astrium Germany, Astrium UK, Galileo Avionica, Zarm The principle feasibility of the HYPER mission has been assessed in the ESA internal CDF study, which is also the starting point for this industrial feasibility study.

6 The HYPER Feasibility Study Main Study Activities Central to this study is the demonstration of system performance by detailed design and analysis, supported by a simulation in particular for drag-free control, precision pointing and thermo-elastic stability of payload elements. Further, a road map for HYPER, taking into account on-going developments in related projects will be defined. The state of art for key components are being assessed and necessary upgrades to meet HYPER needs are identified. Specific ground verification needs will be assessed. Only the measurement of the Lense-Thirring effect and the potential for using the atomic gyroscope as a precision AOCS sensor are setting the requirements for this study. The Atomic Sagnac Unit (ASU) itself, a Mach-Zehnder type interferometer, is the core instrument for both study targets. (Complemented by an integrated Ramsey-Bordé interferometer for the other science objectives.) The detailed design and engineering for the interferometer itself as well as for the supporting laser bench is not subject of this study.

7 The HYPER Feasibility Study Study team and task allocation Customer : Science team: ESA/ESTEC (G.Bagnasco, P. Airey) Consulting (E.Rasel, A. Landragin, P. Bouyer, M.Caldwell) Industrial team: Astrium Germany Astrium UK Zarm Galileo Avionica (Alta Prime, system performance and engineering, AOCS simulator and design, subsystems, AIVT, roadmap, Smart2 LTP optical bench heritage (U. Johann, W. Fichter, L. Szerdahelyi, H. Stockburger, B. Schürenberg) AOCS environment and disturbance; orbit selection; Step, Smart2 platform consulting (S. Kemble, P.Chapman, N. Dunbar) AOCS estimator and DFC sensor assessment (S.Theil, A.Schleicher, Silvia Scheithauer) Optical payload engineering (OB and PST) (G. Cherubini, S. Becucci, A. Romoli) FEEP consulting)

8 The HYPER Feasibility Study Study logic flow Review of present baseline mission and engineering concepts Analysis of SOW payload requirements Environmental and internal disturbance sources characterization Identification of critical areas and design Clarification of open issues and possible trade-off envelopes Orbit selection and ti i ti Reference mission, system concept, architecture and technology definition System performance analysis and simulator development Optical payload engineering Critical subsystems assessment Assessment ASU as AOCS sensor Satellite configuration and subsystems re-assessment based on payload analysis and simulation results We are here Updating from related studies (Smart2, Optimization and performance demonstration (re-fined simulation results) Dec AIVT and specific on-ground testing and verification HYPER mission roadmap End of Jan. 2003

9 The HYPER Feasibility Study Study Work Breakdown Structure

10 The Lense Thirring Effect and HYPER HYPER orbit geometry ( begin of study) Sun synchronous Dawn - dusk 700 km circular Main disturbance: Gravity gradients Air drag Thermal radiation N Ω Pole vo y Ωy (ASU 1,2) Ωx (ASU 3,4) Re = 6371 km Inclination Re ey Ro x Period 98.6 min θ Me ex Ω Equator view Ie ωe Sun S

11 pole equator The Lense Thirring Effect and HYPER Ωy parallel earth spin axis in orbit plane Rotation [rate rad/s] Ωx perpendicular earth spin axis in orbit plane time [s] The Lense-Thirring effect as function of time over two orbit periods. The periodic cycle is half the orbit period. The geodetic de Sitter effect is 40 to 80 times bigger, but rotates perpendicular to orbit plane and is constant for circular orbits

12 The Lense Thirring Effect and HYPER The magnitude of the Lense-Thirring effect (in rad/s) with varying orbit hight, latitude for polar orbit, 90 deg inclination deg latitude deg latitude h=500 h=700 h=1000 h= E-14 Max LT effect (rad/s) 2.7E E E E E E E-14 2E Altitude (km)

13 Illustration of the frame dragging effect (fictive black hole of sun mass and angular momentum)

14 The HYPER Measurement Principle and Payload Definition Aacc Detection λ =780 nm pt=2h/ λ MOT vt Ωrot Rb vt Tdrift/2 vl Preparation Tdrift 2L y x Principles and geometry of the basic Atomic Mach-Zehnder Interferometer The device is sensitive to linear accelerations and rotations in the interferometer plane. Two in-plane counter-propagating Mach Zehnder Interferometers are forming one ASU (Atomic Sagnac unit). Their signals are subtracted to extract the rotation rate signal.

15 The HYPER Measurement Principle and Payload Definition Some ASU parameter values assumed Kγ M m (atomic units) Pt A vt ASU Parameter Laser wavelength λ Size of ASU active zone Фrot [rad] Фacc [rad] Rubidium 780 nm rad/m kg kg m/s 56.2 cm m/s 600 mm x 18.7 mm Ω [rad/s] Aacc [m/s 2 ] Cesium 850 nm rad/m kg kg m/s 32.9 cm m/s 600 mm x 11 mm Ω [rad/s] Aacc [m/s 2 ] 2L = 600 mm, Tdrift = 3 s

16 The HYPER Measurement Principle and Payload Definition Equatorial plane frame dragging measured by HYPER The local inertial frame is probed by the ASU rotation rate measurement The frame rotation is referenced to the global frame by distant star pointing measurement local reference: ASU in-plane rotation rate PST line of sight global reference: guide star position ( Xk, Yk, Zk)

17 The HYPER Measurement Principle and Payload Definition ASU 1,2 Aacc(y) Ωrot(z) ASU 3,4 1 Precision Star Tracker View Normal to orbit plane Aacc(z) Ωrot(y) Optical Bench Structure 2 Inertial Sensors Aacc(x,y,z) Ωrot(x,y,z) The HYPER payload optical bench as an assembly of attitude and acceleration sensors. The two ASU groups ASU1,2 and ASU3,4 (counter-propagating Mach Zehnder Interferometers) The Precision Star Tracker (PST) The two Inertial Sensors (IS)

18 HYPER: ESA Feasibility Study The HYPER Measurement Principle and Payload Definition unambiguous range tune ASU phase set point for inspace calibration and to stay within control range despite gravity gradients (e.g. by laser shirp between interaction regions) control range to be maintained by drag-free and attitude control (DFAC) Rotation rate, linear acceleration, gravity gradients,... ASU signal intensity defined as the difference of two counter-propagating Sagnac units. The signals are illustrating coherence range, un-ambiguity range, linear regime and sensitivity (slope). The high frequency signal (red) stands for the nominal high sensitivity mode (3s drift time) and the low frequency signal (blue) for a fast lower sensitivity mode (AOCS sensor mode; 1s drift time).

19 The HYPER Measurement Principle and Payload Definition Disturbances and sensors for science objectives and drag-free and attitude control Attitude projection effects Lense- Thirring ASU IS PST Guide Star De Sitter External noise: Gravity gradients Drag Thermal fields Charging Magnetic field Thruster noise Measures: Rot, Acc within two ortogonal sensitve planes rad/s m/s2 Bandwidth Hz Transfer budgets Measures: Rot, Acc All directions 10-7 rad/ Hz 100 pm/ Hz m/s 2 / Hz Bandwidth low High possible Transfer budgets Measures: Rot Lateral to line of sight 10-7 rad/10hz Bandwidth medium depends on star magnitude Abberations: Refraction Special relativity General reativity (light bending by frame dragging) Star proper motion Coordinate system transformations from moving S/C to star Extended object

20 The HYPER Measurement Principle and Payload Definition Effects on the ASU phase shift low frequency servo loop a GG estimate Internal ASU phase correction knowledge GG estimate r C,A estimate GG r C,A compensation of gravity gradient acceleration gravity gradient acceleration a GG laser frequency phase Drag-free and attitude control disturbance (drag, etc.) spacecraft a, ω, dω/dt Other effects (magnetic, self gravity, thermo-elastic motion) Atom Interferometer The ASU needs to be isolated from external and spacecraft disturbance by the DFAC (drag-free and attitude control) and precision thermal control (supported by design) The ASU signal itself is used in feedback to stay within control range

21 HYPER Technical Requirements Classification Requirements breakdown LT measurement accuracy 2.5e-15 rad/sec ASU measurement accuracy ASU-PST alignment & stability PST measurement accuracy ASU operational envelope Level-0 gravity gradient knowledge PST/OB alignment PST internal bias and low frequency accuracy linear acceleration dynamic range Level-1 geopotential model OB/RRM alignment PST thermal misalignment drag-free point geopotential pos. knowledge geopotential att. knowledge self gravity magnetic environment temperature timing/jitter PST/OB alignment stability OB/RRM alignment stability star catalogue accuracy aberration effects arithmetic electronic and shot noise timing/jitter acceleration from angular motion acceleration from thermo-elastic motion rate dynamic range rotation of rigid body rotation from thermo-elastic motion alignment RRM/FI Level-2 radiation alignment RRM/AA alignment (set of reqs.) grav. grad. knowledge (for compensation)

22 HYPER Technical Requirements Classification Level-0 Requirements (1) Approach: equal distribution between ASU measurement accuracy and PST measurement accuracy ASU-PST alignment stability to be negligible High pass filtering effect due to data processing 1/10 Lense-Thirring frequency

23 HYPER Technical Requirements Classification Level-0 Requirements (2) ASU Measurement Accuracy one sample accuracy rad/sec 0.3 Hz 1/10 Lense-Thirring frequency PST Measurement Accuracy one sample accuracy: rad 10 Hz relative gain according to high pass filter

24 HYPER Technical Requirements Classification Level-0 Requirements (3) ASU-PST Alignment Stability zero frequency: better than 1 as Hz - 5 Hz: rad/ Hz less than Hz: relative gain according to high pass filter

25 HYPER Technical Requirements Classification Level-1 Requirements (1) Most critical ASU requirements (3σ): gravity gradient knowledge: /sec 2!! displacement of mirror origin and optical path between atoms and mirror <40pm over 3 sec magnetic cleanliness requirements Most critical PST requirements (3σ): centroiding error smaller than as Most critical operational envelope requirements (3σ): (technical driver for DFAC) residual acceleration < m/sec 2 residual rate < rad/sec

26 Orbit Selection Trade-off Gravity gradient altitude dependence Gravity gradient in the order of /s/s GG must be modelled, model fidelity depends on altitude altitude 1000 km requires a 20th order GG model on ground (req /s/s) 2nd order model on board for envelope control (req /s/s) 2.60E E-06 Gravity gradient (/s/s) 2.40E E E E E E Orbital altitude (km)

27 Orbit Selection Trade-off Gravity gradient spatial dependence Gravity gradient model fidelity: 2nd order ( model ) w.r.t. 10th order ( truth ) 1.4E E-10 1E-10 8E E-11 4E-11 2E-11 0 Longitude (deg) Lat

28 Orbit Selection Trade-off Further criteria for orbit altitude selection: Orbit Altitude 1000 km LT effect degraded Smaller number of eclipse days reduces LT degradation (due to small LT magnitude) Total dose: no driving impact Rockot launch to 1000km Sun-synchronous orbit feasible Lower orbit: Higher orbit: increased external disturbance, gravity gradient modelling more complex launcher capability limits

29 Drag-free and Attitude Control (2.AOCS) Design Disturbance Rejection Mechanisms for Interferometer Phase slow drifts Acc. noise residual acceleration Drag-Free Control ASU Sampling ASUmeasured acceleration ASU Phase Control Equivalent ASU Disturbance frequencies < control bandwidth (0.01 Hz) Disturbance frequencies > 0.3 Hz (sampling frequency) Disturbance very low frequencies < typically Hz (8 hours)

30 Drag-free and Attitude Control (2.AOCS) Design Drag-Free and Attitude Control Loops GPS Receiver inertial position Gravity Gradient Estimate Estimated gravity gradient To ASU for compensation purposes 3-axis attitude Star Tracker x-axis attitude X-Axis Attitude Controller Payload sensors PST DFS 1 DFS 2 y/z-attitude 3-axis linear accelerations Y,Z-Axis Attitude Controller X,Y,Z-Axis Linear Acceleration Controller Thruster Selection FEEPs 4 sets of 4 FEEPs each Actuation signal for each FEEP thruster

31 Drag-free and Attitude Control (2.AOCS) Design Noise rejection Rejection of FEEP Noise and Air Drag Noise Main disturbances: FEEP noise and air-drag Figure includes ASU filter effect FEEP noise to be rejected by DFAC loop ASU sampling Drag noise relatively small FEEP noise to be attenuated at low frequencies Relatively low closed loop bandwidth sufficient (0.01Hz) 10-6 Required FEEPs 10-7 AeroDistMax Acceleration

32 Drag-free and Attitude Control (2.AOCS) Design Micro-propulsion Forces and Torque max thrust typ. < 250 micro-n driven by GG torque CoM-DFP distance 13 3 Thruster Configuration total 16 thrusters sets of 4 thrusters Y Z X 24 Solar Array

33 Drag-free and Attitude Control (2.AOCS) Design Drag-free control aspects summary drag-free requirements are moderate ( m/sec 2 /sqrt(hz)) drag-free control concept relies on low frequency phase control within ASU allows relaxed DFS measurement requirement at low frequency => use existing DFS measurement concept with 2 DFS`s imposes configuration requirements solution with minimum mass and power relatively low noise disturbances air-drag (noise) is relatively small due to 1000 km orbit => relatively low bandwidth of drag-free control sufficient, simplifies controller and estimator design maximum thrust 250 micro-n critical points are: dependence on PST measurement accuracy - still to be finally clarified with simulation availability of FEEPs (GOCE FEEP specification sufficient)

34 Payload Configuration Trade-off Principal considerations: The payload configuration and architecture has a large impact on spacecraft configuration, architecture and on DFAC performance and needs to be defined first. The payload configuration is determined by the arrangement, size and budgets of accommodated sensors and by the stability requirements (internal transfer functions) The payload configuration is determined by the ASU intrinsic requirements, concept, design and budgets Three candidate concepts have been identified, driven by different requirement priorities: 1. Previous design concept (cdf report) PST and mirror groups mounted to a common fiducial block (ULE) 2 Inertial sensors in assymetric arrangement with respect to boresight of PST 2. DFAC sensor envelope concept PST and mirror groups mounted to a common fiducial block (ULE) 3 or 4 inertial sensors defining envelope of DFC area arround ASU s 3. DFAC driven concept (selected as baseline) Two inertial sensors in line with the intersection of ASU planes The telescope is moved to a lateral position 4. Integrated ASU fully symmetric concept Both ASU planes are integrated into one atomic beam assembly Telescope, ASU, 2 inertial sensors are in line

35 Optical Bench Configuration Option 1 Previous CDF baseline Payload Configuration Trade-off IRS axis tilted towards ASU axes (projection effects) Virtual DFC point (AOCS reference) close to ASU plane intersection PST in front or inside fiducial block ASU2 ASU2 IRS1 ASU1 PST PST ASU1 y z x Virtual AOCS Optical Bench Ref. Structure COM IRS2 COM Virtual COM AOCS Ref. IRS1

36 Optical Bench Configuration Option 2 Attractive optional concept studied Payload Configuration Trade-off direct measure of 3D gravity gradients by 3 or 4 IRS direct measure of angular rotations on all axes DFC virtual control point can be located arbitrarily very compact architecture PST and mirror block same fiducial frame requires 3 or 4 IRS units and control and read out electronics IRS3 ASU2 ASU2 PST ASU1 y IRS1 IRS4 z ASU1 x Virtual AOCS Ref. COM PST Optical Bench Structure IRS2 COM IRS1 Virtual AOCS Ref. IRS4 Gravity gradients directly measured by IS s (10-10 level)

37 Payload Configuration Trade-off Optical Bench Configuration Option 3 Baseline concept selected in study ASU2 ASU2 PST PST IRS1 ASU1 Virtual AOCS Ref. IRS2 Virtual AOCS Ref. IRS ASU1 y z x COM Optical Bench Structure COM IRS axis, DFC reference and ASU plane intersection collinear PST plate mounted, centered or using full length of bench Thermo-elastic stability less favourable due to asymmetry

38 Payload Configuration Trade-off Optical Bench Configuration Option 3 Integrated 2-D ASU (advanced payload concept) IRS1 PST Virtual AOCS Ref. COM IRS2 sequential generation of two perpendicular MZ IF planes from one atomic beam (same or two subsequent clouds) highly integrated, symmetric concept requires switching of B-field collinear to active laser beams (sig pol) or B collinear to atomic beam axis (pi pol) sequential detection significant instrument development necessary path towards integrated ASU sensor?? ASUxy y ASU z I F z3 x z Mirror frames ASU Y I F Y 1 counterpropagating atomic beam not shown

39 ASU Conceptual design Assumptions made on ASU internal specifications (impacting optical bench interfaces and accommodation) Atomic cloud: Cs (132.9) 1 µk vtemp = 13.7 mm/s vdrift = 200 = 12 mm = 53 = 95 mm Raman laser: wavelength 852 nm pi pulse few µs diam. (80%) 60 mm intensity TBD W/cm 2 freq. stability TBD Average power <m W Detuning capability to keep ASU in control range: TBD Fiber launcher: pol. maintaining fiber core 5 µm, NA =0.2 collimation optics diam 60 mm, feff = 130 mm QWP folding mirror required Possible fiber launcher geometry to be accommodated on optical bench for each Raman laser beam: 80% central intensity (truncated Gaussian beam) 60 mm diam. QWP 200 mm TBC

40 ASU Conceptual design ASU physical envelopes ASU, Raman laser and atomic beam geometry (TBC) Cs/Rb oven, MOT, beam preparation, detection: 300 x 300 x 250 mm 3 tube 200 mm(d) x 700 mm MOT coupling optics included vacuum housing laser,window, µ-metal shield mirror opt. diam. magnetic guide field solenoid RL1 60 mm RL2 RL3 100 mm 100 mm B 300mm 300mm

41 ASU Conceptual design 100 mm 100 mm Geometry of ASU beams and Mach-Zehnder Interferometers 300mm 300mm RL1 RL2 RL3 11 mm It appears necessary to actively change atomic beam position and attitude and interferometer overlap in space

42 ASU Conceptual design 100 mm 100 mm ASU performance degrading thermo-elastic distortions 300mm 300mm Common mode acceleration Optical path change by differential temperature fluctuations of mirror substrates and support (differential acceleration of surfaces) Optical path change by temperature fluctuations in window (acceleration) Tilt of mirror surface Relative in-plane displacement d of counterpropagating Mach- Zehnder Interferometers Common area Raman laser in-plane tilt and shift Relative in-plane tilt of counterpropagating Mach-Zehnder Interferometers

43 HYPER Payload Conceptual Design Payload module precision star tracker Baseline concept to meet resolution requirement 2.5 mas: Optical configuration Ritchey-Chretien telescope Effective focal length Pupil diameter CCD detector 36 m 190 CCD 36 m effective focal length folded into optical bench dimensions (700 mm) Pixel area 512 x 512 Pixels size 13 µm square Tracking matrix 15 x 15 or 17 x 17 FOV as Pixel FOV as Integration time 100 ms Alternative concepts: Optical wavefront sensing: Hubble fine guidance sensor Optical fringes by co-phasing of four small telescopes Gravity probe B concept (radiometry, pyramid split)

44 HYPER Payload Conceptual Design Payload module (Option 3) PST optomechanics and detection ASU fiber launcher inserts PSTbaffle Inertial sensor ASU beam preparation and detection units ASU drift tube housing

45 HYPER Payload Conceptual Design Payload module (Option 3) Back view Optical bench structure (ULE or Zerodur)

46 HYPER Payload Conceptual Design Payload module complete configuration (Option 3) Incl. Fiber injectors and baffle thermal shield

47 HYPER Payload Conceptual Design Payload module optical bench structure (Option 3) Exploded view of optical bench structure (ULE or Zerodur; integrated by hydroxyl bonding technique) ASU drift tube housing: the optical bench structure itself is the vacuum enclosure for the drift tube; the ASU MOT and detection units are to be mounted to the facets

48 HYPER Payload Conceptual Design Payload module optical bench structure (Option 3) Exploded view of optical bench structure (ULE or Zerodur) Mirror group

49 HYPER Payload Conceptual Design Payload module alternative optical bench configuration (Option 2) Inertial Sensor 2 Inertial Sensor 1 Inertial Sensor 3 Inertial Sensor 4

50 HYPER Payload Conceptual Design Payload module thermal design (Option 3) Thermo-elastic distortions of optical bench (preliminary analysis; to be confirmed)

51 HYPER Payload Conceptual Design Thermal radiation control concept S/C inertially fix attitude -> albedo field rotates at orbit frequency arround S/C axis (z) 6 W heat gradient radiated to OB lateral surface, rotating at orbit frequency Thermal stability of PLM requires active control of PLM thermal background (variabe environment in S/C and external radiation field) Uniform thermal background provided by 8 heater mats attached to S/C cylinder structure. Control capability C Temperature level slightly elevated or lowered wrt. PLM to buffer external variation (value TBD). PLM set point temperature 20 C (TBC). Low heater power required (<20W, TBC) Radiative loss to space via baffle minimised by quarz thermal shield in front of PST and additional heater mat wrapped arround the baffle to compensate variable sink temperature (and to control OB set point temperature and internal stationary gradients) Thermal loss through vacuum vent TBD Detailed thermal budget available after thermo-elastic sensitivity from FE model is determined

52 HYPER Payload Conceptual Design The present PLM configuration has been accommodated on a S/C architecture based on the CDF layout Impact: Increased central cylinder diameter and adapter cone Added struts Octagonal shape of S/C bus has thermal advantages (optional concept) Detailed total mass budget pending freezing of PLM baseline and budgets

53 HYPER Payload Conceptual Design The present PLM is isostatically mounted to the central S/C cylinder Elements: Attachement points on optical bench basic structural plates No stress introduced to optical bench No launch locks required Launch loads 12g along axis and 5g laterally assumed PLM Mass 350 kg assumed Heat intake via struts less than 100 mw total (assuming actively controlled cylinder temperature) Strut elements CRFP (blue); Ti (red) Fixation on optical bench by insert technology (see PM2)

54 Conclusions and outlook Present status and work to be done Already done: Previous work analysed and reference mission for study defined Hyper performance requirements breakdown completed Orbit trade-off completed Disturbance environment analysed Payload configuration trade-off completed and baseline defined DFAC simulator developed and almost completed (detailed verification pending) Optical bench and Precision Star Tracker conceptual design and budgets completed Optical bench finite elements model (thermal, structural) almost completed 2. AOCS elements (Inertial sensor and FEEP s) assessment completed Still to be done: DFAC (2.AOCS) simulation campaign Payload module thermo-elastic distortion analysis and design re-finement Precision Star Tracker detailed design Spacecraft configuration and subsystems update Ground testing assessment Hyper road map (technology developments, payload) Payload detailed design and technology development first priority