Instrumentation Payloads - Summer School Alpbach

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ESA UNCLASSIFIED – For Official Use

Instrumentation – Payloads

ESA Summer School 2014 18 July 2014, Alpbach, Austria

Michael Fehringer with contributions from R.Floberghagen, R. Haagmans, R. Bock and F. Heliere Biomass Project Manager ESA Earth Observation Projects Department

Alpbach Summer School | M. Fehringer | 18/07/2014 | EOP | Slide 2


1. Introduction

2. Requirement management

3. Payloads – examples

4. ground coverage - repeat cycle concept

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Introduction

• You are both mission proposers and mission implementers

• Draw clear distinction between “overall goals (= societal

benefit)”, “mission requirements” and derived “system

implementation requirements”

• ESA terminology

• mission requirements Mission Requirements Document

(formulated by the proposing science community and become

responsibility of ESA science departments)

• system implementation requ. System Requirements Doc.

(responsibility with ESA Project Departments)

• These two document are the basis for the mission

implementation and formulate ESA’s responsibility towards the

European science community and funding member states and

govern the mission implementation with industry.

• The SRD becomes the main contractual doc. vs. industry



Introduction

EXAMPLES of successful mission proposals:

• GOCE: Gravity and Ocean Circulation Explorer

• Overall goal: geodesy, ocean circulation, sea level change, etc.

• Mission requirement:

• geoid at 1-2 cm accuracy, gravity anomaly at 1 mGal (10exp-5 m/s2)

• Translates into system requirement

• gravity gradients in 5 – 100 mHz bandwidth at 100 to 11

mEtvos/sqrtHz noise PSD (1mE =10exp-12/sec2)

• Provide orbit at 1 cm accuracy level

• BIOMASS:

• Overall goal: reduce uncertainties in carbon cycle modeling, deforestation

monitoring (10 Gton carbon/year)

• Mission requirement:

• Provide global biomass map in tons/hectare 2x per year

• Translates into system requirement

• Fly a P-band Synthetic Aperture Radar with well specified radiometric

performance and required global coverage period

Steps for today

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Introduction

• Translation of mission requirements to system implementation

requirements is a difficult step

• Can take many years of preparation

• The gap to bridge can only be done by modeling and

simulations

• Tools and algorithms are mostly not fully developed when

mission is selected and implementation starts

• Requires trust and buildup of mutually relevant expertise on

science, ESA and industry side

• However, contracts with industry for implementation needs to

be based on firm performance requirements

give proper attention to this step



GOCE example for req. translation

• How to go from the 1-2 cm geoid req. to a performance

requirement on an instrument the achievement of which

industry has to demonstrate at the in-orbit review?

• Gravity is a conservative force for which the requirement hold that the

trace of the gradient tensor in vacuum is zero (Laplace condition)

• Gravity gradients are the main output of the GOCE satellite

• The sum of the inline elements of the tensor (=trace) is the noise

• This noise is the “quality criterion” for the GOCE mission

• The success criterion the is formulated for the power spectral density in

the measurement band of the GOCE gradiometer

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GOCE performance criterion



Error budget for gravity gradient trace at 100 mHz

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Introduction to payloads

• In most missions the payload is the most complex part and

responsible for cost and schedule overruns

• Consider to use available technologies or timely technology

pre-developments for critical technologies to make sure these

are ready before you kick-off work of big teams in industry

• Industry develops “recurrent platforms”, i.e. reuse and

standardization of platform equipment where mission specific

payloads can be mounted

• Implementation schemes were payload work is started before

the full mission implementation team gets onboard are being

looked at



Payloads for geophysics missions

EXAMPLES:

• the frequency spectrum from visible to radio frequency

• Examples: Synthetic Aperture Radars, Altimetry

• Gravity

• Gradiometry, satellite to satellite tracking, cold atom

interferometry

• Magnetism

• magnetometers

• spectrometry via in-situ measurements

• In situ measurement techniques

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Payloads for geophysics missions - gravity

• Need to fly low

• Signal decays with 1/r2

• The further away, more detail is lost target becomes point source

• Spatial resolution linked to time resolution (Earth orbit, v about 7 km/s)

• Broad features low frequencies

• Small features high frequencies

• In electronics, To measure low frequencies we need integrator (= low pass

filter)

• Satellite to satellite tracking (GPS instrument, precise orbit)

• To measure high frequencies we need differentiator (= high pass filter)

• Gradiometry (e.g. GOCE gradiometer)

• Geopotential can be represented in terms of frequencies in analogy to Fourier

transform, for 3D spherical shape – expansion into spherical harmonics



Payloads for geophysics missions - gravity

Options for gravity missions

• Gradiometer + high/low satellite-to satellite tracking: GOCE

• excellent sensitive to small spatial scales (high degrees)

• Good sensitivity to medium to low scale features

Mission to go for when static gravity field is the objective

• Ranging between two satellites: GRACE

• Range and relative velocity between two satellites at 220 km

distance are used to derive the gravity field (<10 um distance)

• increase in gravity ahead of the pair, the front satellite speeds up

and the distance between the pair increases, changes smaller

than a um/sec in relative velocity detectable

• Better in lower degrees than GOCE (longer measurement baseline

compared to gradiometer), less sensitive to high degrees

Mission to go for when variations in gravity field is main objective

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Gradiometer – 6 free falling test masses


Gradiometer test mass

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Pt-Rh proof mass of 4x4x1

cm and 320 g mass

Accelerometer cage made of

ULE ceramics with gold

electrodes for 6 DOF control

8 electrode pairs per

sensitive element (for

redundancy reasons)

Proof mass grounded by a

25 mm long 5 micron gold wire

Accelerometer Sensor Heads



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X

ZY

Roll

Yaw

Pitch

Accelerometer Servo-Control Loop

Calibration yields relationship between control voltage and force

(acceleration), incl. non-linearities (2nd and 3rd order)

Non-linearities are physically adjusted

Linear combinations of output from different electrodes yield the

tensor components as well as linear and angular accelerations

Maximum redundancy provided through the use of 8 electrode

pairs per proof mass

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Gradiometer


Payload - Gradiometer

Electrostatic Gravity Gradiometer

3 pairs of servo-controlled

capacitive accelerometers on ultra

stable carbon-carbon compound

structure

0.5 m arm length

Accelerometer sensitivity: 2x10-12

m/sec2 rtHz

Structural stability: 0.2 ppm/K

Temperature stability: 10 mK over

200 sec (actively controlled)

Overall stability: few pm in

bandwidth

Mass 180 kg

Power 100 W

Gradiometer bandwidth: 5 to 100

mHz

Used also as AOCS sensor

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Payload - GPS Instrument SSTI

Satellite to Satellite Tracking

Instrument

Dual frequency L1, L2

12 channel GPS receiver

Real time position and velocity

(3D, 3igma) < 100 m, < 0.3 m/s

1 Hz data rate

Science and real time on board

solution for navigation

Precise orbits after ground

processing at 1 cm level

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Satellite Characteristics

3 axis stabilised, nadir pointing,

aerodynamically shaped satellite

5.3 m long, 1.1 m2 cross section,

Launch mass 1050 kg

drag free attitude control (DFACS)

in flight direction employing a

proportional Xe electric propulsion

system (1:100 000 rejection)

Very rigid structure, no moving

parts

Attitude control by magnetorquers

N2 cold gas thrusters for

gradiometer calibration

Body and wing mounted solar

panels

GaAs triple junction solar cells,

1300 W



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Highest sensitivity

accelerometers in space

CHAMP: ~10-9

ms-2

GRACE: ~10-

10 ms-2

GOCE: ~10-

12 ms-2

Ultra-stable Carbon-

Carbon structure with

superior thermo-elastic

stability properties in the

MBW

~ 1 pm over 200 s

~10 mK over 200 s

Continuous operation of

highly accurate ion thrusters

with high thrust and thrust

gradient demands

Main Technical

Challenges


Non-conservative forces and drag compensation

Common mode accelerations rootPSD In-line components:

14X, 25Y and 36Z

- 25Y and 36Z components vary

with air density, winds and lift

- 14X component is stable due to

drag compensation

Data from December 2010

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Differential accelerations

14X and 25Y are essentially

compliant with 2E-12 m/s2/√Hz

requirement

(100x better than any

accelerometer previously flown in

LEO)

36Z is off by factor ∼2

Differential mode acceleration

rootPSD

In-line components:

14X, 25Y, 36Z



10RE

RE + 450km

RE

3485km 1233km

RE + 110km

RE + 450km

RE

3485km 1233km

RE + 450km

RE

3485km 1233km

RE + 450km

RE

3485km 1233km

RE + 110km

10RE

RE + 450km

RE

3485km 1233km

RE + 110km

Magnetic Field – Example Swarm mission

RE = Earth radius ~ 6371km

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Single satellite

Magnetic field magnitude and vector

components

Electric field vector components

Electron density, Ion/Electron Temp.

Air drag

Position, attitude and time

Constellation 3 satellites:

2 side-by-side in low orbit

1 in higher orbit

three orbital planes with two different near-polar inclinations (global coverage)

Launch 2013: 4 years operations

Mission Requirements

accurate enough at satellite altitude to measure the most demanding signals at finest spatial and fastest required temporal sampling



1. Payload Instruments

a. Absolute Scalar Magnetometer

(ASM)

b. Vector Magnetometer (VFM)

c. Electric Field Instrument (EFI)

d. Accelerometer (ACC)

Ref: http://www.esa.int/Swarm

Swarm Mission

Orbits Swarm A & C

a= 462 km, i = 87.35°

ΔRAANA-C = 1.4°

Orbit Swarm B

a= 510 km, i = 87.75°

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System design drivers / considerations:

a. TII Sensors have to placed at satellite front (ram) surface.

b. Minimise spacecraft charging by "positive grounding”, i.e. primary bus positive

pole connected to ground. This concept allows repelling the electrons from the

satellite surface and avoid a satellite charging.

Performances (specified for Swarm mission)

The vector electric field components shall be determined with a random error better

than 5mV/m.

The measurement accuracy of the plasma density shall be better than 1% for densities

greater than 3x109 m-3.

The air drag acceleration vector components shall be determined with a random error

better than 5*10-8 ms-2 in each direction.

The ion and electron temperature shall be determined with an accuracy better than 1%

for densities greater than 1010 m-3.

Electric Field Instrument (EFI)



Configuration & Performance Requirements

DC Mag Random Error at ASMS < 0.3 nT

DC Mag Field at EFI< 10 uT

DC Mag Random Errorat VFMS < 1.0 nT

Random Error of Drag Acceleration Vector

< 5*10 m/sec-8 2

Potential ± 1V

S/C - Plasma<

ACC - STRS Alignment Stability< 0.1 deg (5.1 m)

ASMS - STRS Alignment Stability< 25 arcsec (2.0 m) (goal)

VFMS - STRS Alignment Stability< 1 arcsec (0.5 m)

EFI - STRS Alignment Stability< 0.1 deg (7.2 m)

2555

9060

4945

1970

Flight Direction

ACC - CoM Offset15 mm (± 10 mm (2 )σ)

500kg incl. 99kg fuel; ~1.0 m² cross section 4 years lifetime

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Instrument Accommodation: ASM & VFM/STR

Including Thermal Cover

CFRP Tube

SiC Cube

STR Inner Baffles

VFM



Optical bench: VFM – STR pre-flight alignment

South-east Spain: Calar Alto alignment campaigns

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Absolute Scalar Magnetometer (ASM)

Objective

To perform absolute measurements of the magnetic field magnitude with high accuracy.

To provide the absolute reference for in-flight calibration of the vector magnetometer (VFM).

Sensing Principle

The instrument makes use of the Zeeman effect, which splits the emission and absorption lines of atoms in an ambient magnetic field, respectively. It uses a HF discharge within a gas cell to excite 4He atoms from the 11S0 ground state to the metastable 23S1 state. This metastable level is split by the Earth magnetic field into 3 Zeeman sublevels. The separation of those sublevels is directly proportional to the ambient field strength (eB/2m with m-electron mass).



Performances: Determination of Magnetic field magnitude with

– Resolution : ~1.5 pT/√Hz (from DC to 300Hz, and field range [5 –70µT]),

– Precision (after corrections) : < 0.1 nT

Instrument Budget (for 2 instruments)

– Mass: 2x 3.64 kg (2 DPUs) 1.45 kg (1 Sensor assembly) 2x 1.2 kg (2 sets of harness)

– Power consumption: 9.5 W

– Dimensions: 300 x 248 x 72 mm (DPU)

295 x 136 x 82 mm (Sensor assembly)

– Data rate: 97 bytes/sec

Absolute Scalar Magnetometer (ASM)

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Vector Field Magnetometer (VFM)

Objective

To perform measurements of the Earth's magnetic field vector components with high precision and

Sensing Principle: flux gate magnetometer



Flux gate magnetometer

Fluxgate sensors are typically ring cores of a highly magnetically permeable alloy around which are wrapped two coil windings: the drive winding and the sense winding, consider two halfs (blue and green) AC current applied to drive core into saturation When external field is zero no net flux in sensing coil no signal in sensing coil Source: Imperial College

With external field: one half comes out of saturation earlier than other pickup in sense coil prop. to external field at 2x drive frequency

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Instrument Budget

– Mass: 0.75 kg (DPU) 0.28 kg (Sensor) 0.5 kg (harness)

– Power consumption: 1 W

– Dimensions: 100 x 100 x 60 mm (DPU)

Ø 80 mm (Sensor)


The VFM has been

designed, developed and

manufactured by the

Technical University of

Denmark (DTU).

Performances: determination of magnetic field

vector with

– In the range ±65.000 nT (Earth magnetic field)

– Precision 50 pT rms



System design drivers / considerations:

– Magnetic sensitive instrument => to be located sufficiently far away

from the electromagnetic “dirty” equipment on the satellite bus => on a

deployable boom.

– Sensor axes orientation need to be determined. Co-locate the VFM

sensor with star trackers on a rigid structure (called “optical bench”):


VFM sensor

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ESA UNCLASSIFIED – For Official Use 41


1. Objective:

Characterize the electric field about the Earth by measuring the plasma density, drift, and

acceleration at high resolution.

The EFI Instrument is comprised of two main

sensors: the Thermal Ion Imager (TII) and a set

of two Langmuir Probes (LP). An Electronics

Assembly contains all of the electronics

necessary to control the sensors and contains a

power supply and units for communications with

the Swarm spacecraft.

The Electronics Assembly and TII sensors will

be positioned on the ram face of each Swarm

spacecraft along with the Langmuir probes

positioned on the nadir face of each spacecraft’s




Sensing principle (TII Sensors): Ions enter a narrow aperture slit and are then deflected by a pair of hemispherical grids that create a region having electric fields directed radially inward. Incoming low-energy positive ions are accelerated toward the center of the spherical system, whereas ions with larger kinetic energies travel farther toward the edge of the detector, creating an energy spectrum as a function of detector radius. The resulting image from each TII sensor is a 2-D cut through the ion distribution function, from which one can calculate ion density, drift velocity (2-D), temperature, and higher-order moments.

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1. Sensing principle (Langmuir Probes):

The set of Langmuir Probes provides measurement of electron density,

electron temperature and spacecraft potential.

A bias voltage is applied to the probe and the resulting current, which is

proportional to the plasma charge density, is measured.

To enable simultaneous measurements of electrons and ions, dual probes

are used with one probe biased at a positive potential and the other at a

negative potential.

Instrument Budget:

– Mass: 6.1 kg

– Power consumption: 9.5 W

– Dimensions: 360 x 279 x 210 (TII & electronics)

– Ø 75 x 114 (Langmuir Probes)





1. Objective:

a. Measure the satellite non-

gravitational accelerations at

the satellite orbit (air drag and

solar wind forces).

b. Derive air density models.

Accelerometer (ACC)

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Radar – RAdiation Detection And Ranging

• Fundamentally different to optical instruments

• Time is the essential parameter and the diffraction limit the game

changer d @ l/D

• For geophysics applications 3 radar modes are typically used (e.g. on

Cryosat mission to measure mass and thickness fluctuations of land and

marine ice fields)

• altimetry over “flat” areas like oceans and central ice caps

• SAR – synthetic aperture radar

• Imaging radar with improved spatial resolution

• Interferometric SAR

• Interferometry either with repeating orbits (coming back to

same scene after defined duration) or simultaneously with

two antennas

• Resolves range ambiguities at slopes in terrain that can

occur in normal SAR



SAR – synthetic aperture radar

• Optical systems

• Have excellent angular resolution d @ l/D

• Results (images) are “easy” (intuitively) to interpret

• Cannot image through clouds, need sunlight

• Not suitable for interferometry in geophysical applications

(wavelength too short)

• SAR systems

• Radars have bad angular resolution d = l/D

• Data are not intuitively understood

• No rotational symmetry, azimuth and range treated differently

• Need massive and complex processing

• Penetrate clouds, no need for daylight

• Well suitable for interferometry in geophysics applications

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Synthetic Aperture Radar

Synthetic Aperture Radar : imaging radar mounted on a moving platform.

Real Aperture Radar

Poor Azimuth Resolution #km

SAR

Azimuth resolution independent of frequency and distance

Prop l



SAR Geometry and range resolution

T: pulse length

Chirp signal (frequency modulated pulsed waveforms) used to improve range resolution.

c0 : speed of light Br: Chirp bandwidth

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SAR processing steps

1. Convolution of the raw data with the range reference function (Chirp). 2. Convolution with the azimuth reference function, which changes from near

to far range.



Differential SAR interferometry Example : Subsidence detection

Zoom over the city

Estimated subsidence over Mexico City obtained with two TerraSAR-X images acquired with a 6-month difference (overlay of reflectivity and phase).

Mean deformation velocity estimated over Mexico City

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SAR frequency and applications

20 m resolution, C-band, radar illumination from the left

1 m resolution, X-band, radar illumination from the right

Pyramids of Giza, Egypt.



0

100

200

300

400

500

600

L band image Biomass map from P-band

Tropical forest (French Guiana)

1

1

2

L-band

• Forest regrowth area, AGB=180 t/ha, (1)

NOT distinguished from neighbouring

intact forest with AGB>400 t/ha (2).

• Young and sparse plantations, AGB < 10

ton/ha (3) distinguished from bare soil

(4)

4

P-band

• Forest regrowth area, AGB=180 t/ha, (1)

distinguished from neighbouring intact

forest with AGB>400 t/ha (2).

• Young and sparse plantations, AGB < 10

ton/ha (3) NOT distinguished from bare

soil (4)

3

Synergetic use of P- and L-band SAR data

1

1

2

4 3

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A single P-band satellite can deliver 3 independent types of information for biomass

PolSAR (SAR Polarimetry)

x

y

z

o

PolInSAR (Polarimetric SAR Interferometry)

x

y

z

o

Height

Tomo SAR (SAR Tomography)

x

y

z

o

Height

EUSAR, 05 June 2014 Page 53



Major SAR Requirements – example Biomass

Parameter Requirement Instrument type P-band full polarimetric SAR

Centre frequency 435 MHz (P-band)

Bandwidth 6 MHz (ITU allocation)

Incidence angle (near) Threshold: 23; Target: 25

Cross-polarisation ratio ≤–25 dB (threshold); ≤ -30 dB (goal)

Spatial res. ( 6 looks) 60 m (across-track) 50 m (along-track)

Noise equivalent 0 Threshold: –27 dB; Target: –30 dB

Total ambiguity ratio 18 dB

Radiometric stability 0.5 dB RMS

Abs. radiometric bias 1.0 dB

Dynamic range 35 dB

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Cryosat

measure mass and thickness

fluctuations of land and marine ice fields

• SAR/interferometric radar altimeter

• Ku-band (13.7 GHz)

• 717 km, non sun synchronous, 0.25

deg nodal regression/day

• 369 days repeat cycle, 30 d subcycle

• 0.1 deg pointing error

• 0.001 deg pointing stability

• 670 kg

• 1600 W

• 320 Gbit/day

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Sentinel -1

Land and ocean monitoring

• C-band SAR (5.4 GHz)

• 4.8 kW, up to 1.2 TB/day

• 12 m antenna

• 2300 kg

• Up to 400 km swath width

• Down to 5x5 m resolution

• 6 days revisit with 2 satellites

• Attitude accuracy and knowledge

<0.01 deg/axis and <0.003 deg/axis

• 10 m position knowledge (3 sigma)



In-situ measurements – example Rosetta lander

1. APXS: Alpha Proton X-ray Spectrometer (chemical composition)

2. CONSERT: COmet Nucleus Sounding Experiment by Radiowave

Transmission (studying the internal structure of the comet nucleus with

Rosetta orbiter)

3. COSAC: The COmetary SAmpling and Composition (detecting and

identifying complex organic molecules) – gas analyser

4. PTOLEMY: Determining and Understanding Light elements to

understand the geochemistry of light elements, such as hydrogen,

carbon, nitrogen and oxygen – gas chromatography/mass

spectrometer

5. MUPUS: MUlti-PUrpose Sensors for Surface and Sub-Surface Science

(studying the properties of the comet surface and immediate sub-

surface)

6. SD2: Sampling, drilling and distribution subsystem (drilling up to 23

cm depth and delivering material to onboard instruments for analysis)

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In-situ measurements – example Rosetta lander

1. MUPUS: MUlti-PUrpose Sensors for Surface and Sub-Surface Science

(studying the properties of the comet surface and immediate sub-

surface)

2. SD2: Sampling, drilling and distribution subsystem (drilling up to 23

cm depth and delivering material to onboard instruments for analysis)

3. SESAME: Surface Electric Sounding and Acoustic Monitoring

Experiment (probing the mechanical and electrical parameters of the

comet), comprising: CASSE (Comet Acoustic Surface Sounding

Experiment), DIM (Dust Impact Monitor), and PP (Permittivity Probe).



Ground coverage – repeat cycle concept

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Coverage build-up

INT phase with 4 days repeat cycle

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