ep test experiment environment - oca

EP test experiment Environment – 29-30/01/2013 Microscope colloquium II

29/01/2013

A. ROBERT DCT/PO/EU

EP test experiment Environment

EP test experiment Environment – 29-30/01/2013 Microscope colloquium II2

Contents

1. Thermal environment

2. Magnetic cleanliness

3. Satellite self gravity

4. Micro perturbations



Justification

� Instrument bias thermal sensitivity and scale factor thermal sensitivity.� Compromise between satellite thermal requirements and instrument sensitivity.� Major contributor in the mission error budget

Main requirements

� SU� On SU skin: 1 mK @fep� Thermal axial gradient on SU skin: 2.5 mK/m @fep

� FEEU� 5mK @fep at the interface inertiel mode� 3mK @fep at the interface spin mode

� ICU� 1K @fep at the interface

SU1

SU2

SUMI

FEEU

ICUME



Thermo-mechanical architecture

� Driven by � SU accommodation constraints

� Satellite MCI budget goals� Instrument thermal stability requirements

� Concept� SU in the platform near the CDG of the satellite� Structure supporting the SU and the FEEU: the payload assembly subsystem: PAS

» Including the thermal control hardware» Allowing the centering of proof masses» Offering a high thermal stability to SU and FEEU at fep» Guaranteeing the thermal control of FEEU (6 W power dissipation of each)

� PAS is mounted on the –X panel (anti solar panel)

ICUME

Battery

RW

OBC GNSS

PCDU

CGPSTank

ZSAT

YSAT

XSAT

PayloadModule



PAS design

� A two-stage structure� First stage support the FEEU and its radiator

» Mechanically linked and thermally insulated from-X panel by six titanium blades

� Second stage support the SU and magnetic shielding» Mechanically linked and thermally insulated from

from the first stage by six titanium blades» FEEU radiator protected from external IR Earth fluxes

by a specific baffle

� Each stage is:� Covered by MLI� Conductively decoupled from the satellite

PAS represents a thermal cavity insulated from the satellitewith its autonomous passive thermal control

SU

FEEU1

FEEU2

FEEU radiator

I/F with-Xsat panel

SU stage

FEEU stage

SU shielding

Titanium bladeSU stage

Titanium blade FEEU stage



Performances

� Verified by thermal analysis and qualification test performed in 2009

� Only one requirement is not compliant» FEEU stability: 5mK @fep at the interface inertial mode» Between 5.5mK and 8.8mK depending on the date and the proximity of the eclipse period

� This was taken into account in the mission error budget previously presented



Justification

� The magnetic disturbance acceleration is given by the relation:

where and are the magnetic susceptibility and density

of the proof mass and B the magnetic field.

� A differential disturbance acceleration could be induced by:» The using of platinum-rhodium and titanium for the proof masses» The magnetic environment of the instrument (earth and satellite)

Requirements of the magnetic environment

� Magnetic cleanliness is not clearly specified because there is no analytical

expression

� So only the effects on the instrument are specified:

∇=

2

m0

m B.2

m ρµχΓ

mχ mρ

Parameter Specification Domain MIS-250 DC effect 2.5 10-9 ms-2 On each inertial sensor mass MIS-251 Stochastic

variations 4.10-14 ms-2/√Hz On each inertial sensor

MIS-252 Tone error at fep

2.10-16 ms-2

10-12 ms-2 Differential effect on the two masses Common mode acceleration on both masses, to be included in the drag-free residue

MIS-253 Tone error at 2 fep

4.10-15 ms-2

2.10-11 ms-2 Differential effect on the two masses Common mode acceleration on both masses, to be included in the drag-free residue



Analysis

� A two steps analysis� Instrument mesh modeling to use a finite elements simulations tool (Flux 3D)

» Compute of the magnetic field and its gradient in each proof mass in response to a field (earth) or a magnetic dipole (satellite) calibrated.

� Estimation of the differential acceleration perturbation and its spectral composition

(DC, fep, 2fep) in inertial or spin mode» Using the knowledge of the in flight magnetic environment (earth and satellite hypothesis)

and the sensitivity calculated at the first step.

� Geometric definition and mesh of the SU (Flux 3D)» Modeling the magnetic parts

of the instrument: invar» The proof mass

Sole plate

External casingHat



External shielding

� Error budget� Tone error at fep: 50 times the requirement.

� Need of an external shielding� Made of high permeability material (Supranhyster 50)

� To strongly reduce the residual field inside the instrument

� Geometric definition of the shielding



Performances

� The error budget complies with a big margin

Complementary analysis

� Shielding effect will be tested on the FM in 2013

� Magnetic momentum of the satellite equipments are going to be measured

in order to confirm magnetic satellite hypothesis» SST, PCDU, Battery, OBC …



Justification each satellite element induces a gravity acceleration applied to the test masses

� Self gravity� Specification DC value linked to the instrument full scale (2.5 10-7 m/s2)

� Gravity gradient� Test mass centring error: requirement of 20µm � Limit the variation of the gravity gradient due to a motion of a mass on board the

satellite (thermoelastic deformation) in order to minimize the differential acceleration

error

Main requirementsREF Parameter Specification Domain

MIS-910 2 10-9s-2Hz-1/2 about fep

MIS-920 5 10-12 s-2

Tone error at fep

MIS-930 5 10-10 s-2

Tone error at 2 fep

MIS-940

S/C gradient of gravity components

2.5 10-7 s-2 DC Each axis

MIS-950 Spacecraft self-gravity

X: 5 10-9 ms-2 Y,Z: 2 10-8 ms-2

X: 10-12 ms-2 Y,Z: 10-12 ms-2

X: 2.10-11 ms-2 Y,Z: 2.10-11 ms-2

DC At fep, to be included in the drag-free residue At 2fep, to be included in the drag-free residue



Analysis

� Use of the satellite finite element model» 135,000 nodes

� Mass discretization, NASTRAN elements are

replaced by concentrated masses at their centre» Computation of satellite self-gravity

� Determination of the nodes displacements underthermal loading (provided by thermal analysis)

» Computation of the fep gravity gradient variation at test masses locations

Performances

� Compliant with the required values



Self gravity and cold gas propulsion system

� The motion of the gravity centres of the gas mass in the 6 tanks has been

evaluated considering:� Gas consumption

� Variation of the local gravity at tank level

� Centrifugal force in rotating mode� Variation of the temperature of the tanks

� Gas convection

� Compliance has been demonstrated



Overview

� Micro-perturbations are all physical non periodic phenomena that can temporarily disturb measurements

� List of possible micro-perturbations : � Micro gliding, local buckling, Laplace and Lorentz forces, antenna emission forces, meteorites

and debris effects, mechanical relays, wheels movement, moving mass in the thrusters, liquid sloshing, MLI thermoelastic clanks, tanks release effects …

� Effect on the experiment� Saturation of the instrument involving not available measurements

» Deterioration in the efficiency of data processing

� Possibility of loss of the control loop of the instrument during a measurement session

� The control plan is articulated around the following themes� Design rules (Laplace electromagnetic forces, no mechanism, passive thermal control)

� Dimensioning to prevent micro-gliding at the junctions and local buckling in the structures

induced by thermo-elastic distortion.

� Statistical analysis of the debris and meteorites and the effects on the instrument, using the ESA

software MASTER (Meteoroid and Space Debris Terrestrial Environment Reference). This

software provides space debris flux and spatial densities.

� Tests (choice of external MLI material, tank pressure release, effect of the moving mass in the

thrusters)



Tests performed

� On the thruster� To estimate the disturbance induced by moving mass

» The gas flow is controlled via a piezo control loop:� A mass of about 5g moves to control the gas flow and

generates reaction forces on the satellite

� The test consists in measuring the disturbances caused by a thrust step of 2µn (representative flight conditions given by AACS)

� Compliant with the instrument constraints

� On the CGPS tank� The tank consist of a metallic liner overwrapped with a high-strength fiber/polymeric

matrix resin composite whose size increases with pressure (350 bars, 5mm on the length). In flight some suggested the possibility of clank during depressurization.



� Test objective» Measuring the tank deformation to estimate the moving mass during the clank (local and

global deformation)

� Principle» Digital image correlation for local deformation» Capacitive measurement for global deformation

� Results» No global deformation» Some local deformations whose effects are consistent with mission performances

requirements



� To choice the MLI material� Test objective

» To provide experimental data leading to the choice of the optimal MLI with respect to the minimization of micro-perturbations induced by thermally loaded external MLI

» Eight MLI samples have been tested during two tests

� Test setup» Measurements performed during thermal cycling in vacuum using the test facilities in the IABG

Space Test Centre» The MLI samples were fixed to sandwich composite panel (Demeter)» The panel response to MLI clank was monitored by 4 high sensitive accelerometers

� Results» Quantitative evaluation on the number of thermally

induced MLI micro-vibration events and choice of the MLI» Quantitative analysis and the assessment of the effects

on the mission performances by:� Using a dynamic model of the panel and � Modeling the clank itself

» Conclusion: effects are consistent with the mission

ep test experiment environment - oca

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