status of the esa feasibility study hypersci2.esa.int/hyper/hypersymposium1/hypercd/hyper symp-...

Status of the ESA Feasibility Study

HYPER

Ulrich Johann

Astrium GmbH

(for the industrial team)

CNES, Paris

[email protected]

HYPER: ESA Feasibility StudyHYPER Symposium 1Fundamental Physics and Applications in SpaceCNES, Paris, 5.11.2002

Presentation outline:Scientific Objective of the HYPER MissionThe present HYPER Mission Baseline The HYPER Feasibility Study

The Lense-Thirring Effect and HYPERHYPER Measurement Principle and Payload Definition

HYPER Technical Requirements Classification Orbit Selection Trade-offPayload 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)© Astrium

HYPER: ESA Feasibility StudyHYPER Symposium 1Fundamental Physics and Applications in SpaceCNES, Paris, 5.11.2002 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.

© Astrium

HYPER: ESA Feasibility StudyHYPER Symposium 1Fundamental Physics and Applications in SpaceCNES, Paris, 5.11.2002 HYPER Mission Baseline; June 2002

767.4 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 m2 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

Spacecraft launch mass:

AOCS:

Propulsion:

Power:

Telemetry:Ground segment:

Dawn-dusk, Sun-synchronous orbit at 700 km altitude, 98.2° inclination, 98.6 minLow-cost Rockot launch vehicle from Plesetsk CosmodromeTo be based on road map results (in terms of technical readiness, 2009)2 years (nominal)

Orbit:Launcher:Launch year:Mission lifetime:

247.5 kg203 W937 mm diameter x 1300 mm height

Optical elements for coherent atom manipulationHigh-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 atomsHigh-precision µW synthesiser for the hyperfine transitions of caesium or rubidium

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)

HYPER: ESA Feasibility StudyHYPER Symposium 1Fundamental Physics and Applications in SpaceCNES, Paris, 5.11.2002 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.

© Astrium


Main Study Activities

The HYPER Feasibility Study

•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-Bordéinterferometer 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.

© Astrium


Study team and task allocation


Customer : ESA/ESTEC (G.Bagnasco, P. Airey)

Science team: Consulting (E.Rasel, A. Landragin, P. Bouyer, M.Caldwell)

Industrial team:Astrium Germany 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)

Astrium UK AOCS environment and disturbance; orbit selection; Step, Smart2 platform consulting(S. Kemble, P.Chapman, N. Dunbar)

Zarm AOCS estimator and DFC sensor assessment(S.Theil, A.Schleicher, Silvia Scheithauer)

Galileo Avionica Optical payload engineering (OB and PST) (G. Cherubini, S. Becucci, A. Romoli)

(Alta FEEP consulting)

© Astrium


Analysis of SOW

payload requirements Review of present baseline

mission and engineering concepts

Environmental and internal disturbance sources characterization

Identification of critical areas and design

Clarification of open issues and possible trade-off envelopes

Reference mission, system concept , architecture and technology definition

System performance analysis and simulator

development

Optical payload

engineering

Critical subsystems assessment

Satellite configuration and subsystemsre-assessment based on payload analysis

and simulation results

Updating from related studies (Smart2,

Optimization and performance demonstration (re-fined simulation results)

Assessment ASU as

AOCS sensor

AIVT and specific on-ground testing and verification

Orbit selection and ti i ti

HYPER mission roadmap

Study logic flow

Dec. 2002

End of Jan. 2003


6.6.2002

We are here

© Astrium


Study Work Breakdown Structure


© Astrium

HYPER: ESA Feasibility StudyHYPER Symposium 1Fundamental Physics and Applications in SpaceCNES, Paris, 5.11.2002 The Lense Thirring Effect and HYPER

N

Ωx (ASU 3,4)

θ

Reey

ex

Ie ωe

Rox

yvo Ωy (ASU 1,2)

Ω EquatorMe

S

Ω Pole

Sun

Main disturbance:•Gravity gradients•Air drag•Thermal radiation

HYPER orbit geometry( begin of study)

•Sun synchronous•Dawn - dusk•700 km circular •Re = 6371 km•Inclination 90° + 8.2°•Period 98.6 min

view

© Astrium


equatorpoleR

otat

ion

[rate

rad/

s]

time [s]

Ωy parallel earth spin axisin orbit plane

Ωxperpendicular earth spin axis in orbit plane

© Astrium

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


The magnitude of the Lense-Thirring effect (in 10-14 rad/s) with varying orbit hight, latitude for polar orbit, 90 deg inclination

0

0.5

1

1.5

2

2.5

h=500

h=700

h=1000

h=1200

90 deg latitude

0 deg latitude

2E-142.1E-14

2.2E-142.3E-14

2.4E-142.5E-14

2.6E-142.7E-14

2.8E-14

500 600 700 800 900 1000 1100 1200 1300

Altitude (km)

Max

LT

effe

ct (r

ad/s

)

© Astrium

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

© Astrium


y

x

2L

RbΩrot

Aacc

λ =780 nm

Tdrift

vT

vL

pT=2h/ λ

vT·Tdrift/2

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.

Detection

MOT

The HYPER Measurement Principle and Payload Definition

Preparation

© Astrium



Some ASU parameter values assumed

6.65·107·Aacc [m/s2]7.250·107·Aacc [m/s2]Фacc [rad]1.331·107· Ω [rad/s]1.450·107· Ω [rad/s]Фrot [rad]

600 mm x 11 mm600 mm x 18.7 mmSize of ASU active zone

0.0073 m/s0.0125 m/svt32.9 cm256.2 cm2A1.56·10-27 kg·m/s1.7·10-27 kg·m/sPt13385m (atomic units)2.13·10-25 kg1.36·10-25 kgM7.392·106 rad/m8.055·106 rad/mKγ850 nm780 nmLaser wavelength λ

CesiumRubidiumASU Parameter

2L = 600 mm, Tdrift = 3 s

© Astrium

Xk Yk, Zk,( )

PST line of sight

local reference: ASU in-plane rotation rate

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



global reference: guide star position

© Astrium



Aacc(z)

Ωrot(z)

Ωrot(y)

Aacc(y)

1 Precision Star Tracker

Optical Bench Structure

ASU 1,2

ASU 3,4View

Normal to orbit plane

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)

© Astrium


unambiguous range


tune ASU phase set point for in-space 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).

© Astrium


Disturbances and sensors for science objectives and drag-free and attitude control


Attitude projection effects

ASU PSTISLense-

Thirring Guide Star

External noise:•Gravity gradients

•Drag

•Thermal fields

•Charging

•Magnetic field

•Thruster noise

Measures:

Rot, Acc within two ortogonal sensitve planes10-12 rad/s10-13 m/s2

Bandwidth

0. 001 -0.3 Hz

Measures:Rot, AccAll directions10-7 rad/√Hz100 pm/√Hz10-14 m/s2/√HzBandwidth lowHigh possible

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

Transfer budgets Transfer budgetsDe Sitter

© Astrium



AtomInterfero-

meter

InternalASU phasecorrection

compensationof gravity gradient

acceleration

gravity gradientacceleration

Drag-free andattitude control

spacecraft

GG estimate

GG

rC,A estimate

rC,A

Other effects(magnetic, self gravity,thermo-elastic motion)

a, ω, dω/dt

aGG

phaselaser

frequency

disturbance(drag, etc.)

aGG estimate

low frequency “servo” loopEffects on the ASU phase shift

knowledge 10-11

•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© Astrium


geopotentialmodel

geopotentialpos. knowledge

geopotentialatt. knowledge

self gravity

gravity gradientknowledge

magneticenvironment

temperature

timing/jitter

radiation

alignment(set of reqs.)

ASUmeasurement

accuracy

PST/OBalignment

OB/RRMalignment

PST/OBalignmentstability

OB/RRMalignmentstability

ASU-PSTalignment& stability

PST internal bias andlow frequency accuracy

PST thermalmisalignment

star catalogueaccuracy

aberration effects

arithmetic

electronicand shot noise

timing/jitter

PSTmeasurement

accuracy

acceleration@ drag-free point

acceleration fromangular motion

acceleration fromthermo-elastic motion

linear accelerationdynamic range

rotation ofrigid body

rotation fromthermo-elastic motion

ratedynamic range

alignmentRRM/FI

alignmentRRM/AA

grav. grad. knowledge(for compensation)

ASUoperationalenvelope

LT measurementaccuracy

2.5e-15 rad/sec

LevelLevel--00

HYPER Technical Requirements Classification

Requirements breakdown

LevelLevel--11

LevelLevel--22

© Astrium



Level-0 Requirements (1)

• Approach: “equal” distribution between ASU measurement accuracy and PST measurement accuracy

• ASU-PST alignment stability to be negligible

1/10 Lense-Thirring frequency

© Astrium

• “High pass” filtering effect due to data processing




• ASU Measurement Accuracy– one sample accuracy 5.0 10-12 rad/sec (3sig) @ 0.3 Hz

1/10 Lense-Thirring frequency

• PST Measurement Accuracy– one sample accuracy:

1.2 10-8 rad (3sig) @ 10 Hz– relative gain according

to “high pass” filter

© Astrium




• ASU-PST Alignment Stability– zero frequency: better than 1 as– 3.5 10-5 Hz - 5 Hz: 1.75 10-9 rad/√Hz– less than 3.5 10-5 Hz: relative gain according to “high pass” filter

© Astrium




• Most critical ASU requirements (3σ):

– gravity gradient knowledge: 2.4 10-11 1/sec2 !!– 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 2 10-3 as

• Most critical operational envelope requirements (3σ):(technical driver for DFAC)

– residual acceleration < 1.7 10-8 m/sec2

– residual rate <4.3 10-8 rad/sec

© Astrium

HYPER: ESA Feasibility StudyHYPER Symposium 1Fundamental Physics and Applications in SpaceCNES, Paris, 5.11.2002 Orbit Selection Trade-off

Gravity gradient altitude dependence

• Gravity gradient in the order of 1 10-6 1/s/s– GG must be modelled, model fidelity depends on altitude– altitude 1000 km requires a

• 20th order GG model on ground (req. 2.4 10-11 1/s/s)• 2nd order model on board for envelope control (req. 3.4 10-8 1/s/s)

1.90E-06

2.00E-06

2.10E-06

2.20E-06

2.30E-06

2.40E-06

2.50E-06

2.60E-06

6700 6800 6900 7000 7100 7200 7300 7400 7500

Orbital altitude (km)

Grav

ity g

radi

ent (

/s/s

)

© Astrium


Gravity gradient spatial dependence

• Gravity gradient model fidelity: 2nd order (“model”) w.r.t. 10th order (“truth”)

0108

216

324

-80 -4

8 -16 16

4880

02E-114E-116E-118E-111E-101.2E-101.4E-10

Longitude (deg)

Lat

© Astrium


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: increased external disturbance, gravity gradient modelling more complex

Higher orbit: launcher capability limits

© Astrium


Drag-free and Attitude Control (2.AOCS) Design

Disturbance Rejection Mechanisms for Interferometer Phase

slow drifts

ASU-measured

acceleration

© Astrium

ASU Sampling

Acc. noiseASU Phase

ControlDrag-Free

Control

Equivalent Phase @ ASU

residual acceleration

Disturbance rejection @ very low frequencies < typically 3.5 10-5 Hz(8 hours)

Disturbance rejection @ frequencies> 0.3 Hz (sampling frequency)

Disturbance rejection @ frequencies < control bandwidth (0.01 Hz)

© Astrium



• Drag-Free and Attitude Control Loops

inertial position

DFS 1

PST

DFS 2

Payload sensors

GPSReceiver

To ASU for compensation

purposes

Estimatedgravity gradientGravity Gradient

Estimate

StarTracker

ThrusterSelection

Actuation signal for

each FEEP thruster

FEEPs

4 sets of4 FEEPs

each

© Astrium

X,Y,Z-AxisLinear Acceleration

Controller

X-AxisAttitude Controller

Y,Z-AxisAttitude

Controller

3-axis attitude

y/z-attitude

3-axis linear accelerations

x-axis attitude



Noise rejection

10-4 10-3 10-2 10-1 100 10110-12

10-11

10-10

10-9

10-8

10-7

10-6

RequiredFEEPsAeroDistMaxAcceleration

– Drag noise relatively small– FEEP noise to be attenuated at low

frequencies– Relatively low closed loop

bandwidth sufficient (0.01Hz)

FEEP noise to be rejected by DFAC loop• Rejection of FEEP Noise and Air Drag Noise

– Main disturbances: FEEP noise and air-drag– Figure includes ASU filter effect

ASU sampling

© Astrium



Micro-propulsion

21

Solar Array

24

23

22

313

414

122

11

1

YZ X

• Forces and Torque– max thrust typ. < 250 micro-N– driven by

• GG torque• CoM-DFP distance

• Thruster Configuration– total 16 thrusters– 4 sets of 4 thrusters

© Astrium



Drag-free control aspects summary

– drag-free requirements are moderate (≈1 10-8 m/sec2/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)

© Astrium

HYPER: ESA Feasibility StudyHYPER Symposium 1Fundamental Physics and Applications in SpaceCNES, Paris, 5.11.2002 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 PST2. 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´s3. DFAC driven concept (selected as baseline)• Two inertial sensors in line with the intersection of ASU planes• The telescope is moved to a lateral position4. Integrated ASU fully symmetric concept• Both ASU planes are integrated into one atomic beam assembly• Telescope, ASU, 2 inertial sensors are in line

© Astrium


Optical Bench Configuration Option 1

Previous CDF baseline

• 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

y

xzOptical Bench Structure

ASU2

ASU1IRS1

IRS2

PST

COM

Virtual AOCSRef.

COM

ASU2

ASU1PST

IRS1COMVirtual

AOCSRef.

© Astrium


Attractive optional concept studied

• 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


Optical Bench Structure

ASU2

IRS1 IRS2

PST

Virtual AOCSRef.

ASU2

ASU1PST

ASU1 Virtual AOCSRef.

y

xz

IRS3

IRS4

IRS1 IRS4

COM COM

Gravity gradients directly measured by IS´s (10-10 level) © Astrium



Baseline concept selected in study

y

xzOptical Bench Structure

Virtual AOCSRef.

ASU2

ASU1IRS1 IRS2

PST

COM

Virtual AOCSRef.

ASU2

ASU1

PST

IRS

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

© Astrium



Integrated 2-D ASU (advanced payload concept)

© Astrium

ASUxy

IRS1

IRS2

Virtual AOCSRef.

PST

COM

ASUzIFz 3

ASUYIFY 1

yx

Mirror framesz

counterpropagating atomic beam not shown

• 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??

HYPER: ESA Feasibility StudyHYPER Symposium 1Fundamental Physics and Applications in SpaceCNES, Paris, 5.11.2002 ASU Conceptual design

Assumptions made on ASU internal specifications(impacting optical bench interfaces and accommodation)

Raman laser:

wavelength 852 nm pi pulse few µs diam. (80%) 60 mm intensity TBD W/cm2

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 QWPfolding mirror required

Atomic cloud:

Cs (132.9) 1 µKvtemp = 13.7 mm/s vdrift = 200 mm/s @RL1 = 12 mm dia @RL2 = 53 mm @RL3 = 95 mm

Possible fiber launcher geometry to be accommodated on optical bench for each Raman laser beam:

QWP

80% central intensity (truncated Gaussian beam)

60 mm diam.

200 mm TBC © Astrium

© Astrium

ASU physical envelopes•ASU, Raman laser and atomic beam geometry (TBC)

300mm

tube 200 mm(d) x 700 mm

RL1 RL2 RL3

Cs/Rb oven, MOT, beam preparation, detection:

300 x 300 x 250 mm3

MOT coupling optics

included

100 mm

100

mm

vacuum housingµ-metal shieldmagnetic guide field solenoid

B


laser,window, mirror opt. diam. 60 mm

300mm

© Astrium

100 mm

100

mm

300mm 300mm

11 mm


It appears necessary to actively change atomic beam position and attitude and interferometer overlap in space

Geometry of ASU beams and Mach-Zehnder Interferometers

RL2RL1 RL3

100 mm

100

mm

300mm 300mm

Common area

Optical path change by temperature fluctuations in window (acceleration)

Common mode acceleration

Relative in-plane displacement d of counterpropagating Mach-Zehnder Interferometers


ASU performance degrading thermo-elastic distortions

Optical path change by differential temperature fluctuations of mirror substrates and support (differential acceleration of surfaces)

Tilt of mirror surface

Relative in-plane tilt of counterpropagating Mach-Zehnder Interferometers

Raman laser in-plane tilt and shift

© Astrium


HYPER Payload Conceptual Design

Payload module precision star trackerBaseline concept to meet resolution requirement 2.5 mas:

+- 15 as0.074 as100 ms

FOVPixel FOVIntegration time

512 x 51213 µm square 15 x 15 or 17 x 17

CCD detector•Pixel area•Pixels size•Tracking matrix

36 m190

Optical configuration•Ritchey-Chretien telescope•Effective focal length•Pupil diameter

36 m effective focal length folded into optical bench dimensions (700 mm)

CCD

Alternative concepts:Optical wavefront sensing: Hubble fine guidance sensorOptical fringes by co-phasing of four small telescopesGravity probe B concept (radiometry, pyramid split)

© Astrium



Payload module (Option 3)

ASU beam preparation and detection units

PST opto-mechanics and detection

PSTbaffle

ASU drift tube housing

ASU fiber launcher inserts

Inertial sensor

© Astrium



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

© Astrium



Payload module optical bench structure (Option 3)Exploded view of optical bench structure (ULE or Zerodur)

Mirror group

© Astrium



Payload module alternative optical bench configuration (Option 2)

Inertial Sensor 1

Inertial Sensor 2

Inertial Sensor 3

Inertial Sensor 4

© Astrium



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

© Astrium



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 +-0.2 °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

© Astrium



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

© Astrium



The present PLM is isostatically mounted to the central S/C cylinderElements:•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)

© Astrium

© Astrium

HYPER: ESA Feasibility StudyHYPER Symposium 1Fundamental Physics and Applications in SpaceCNES, Paris, 5.11.2002 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

© Astrium

status of the esa feasibility study hypersci2.esa.int/hyper/hypersymposium1/hypercd/hyper symp-...

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