AROSAT Mission and Spacecraft Configuration

AROSAT Mission objectives (1/5)

AROSAT is a satellite system optimized to provide maximum ground resolution for target identification and analyses at the expense of revisiting frequency

AROSAT Mission objectives (2/5)Primary objectives:

A) achieve panchromatic ( PAN) Very High Resolution (<0.5m ) imaging of artifacts, infrastructure, small fixed or mobile objects up to +-35° from nadir;

B) improve image features: high dynamic range (targets with high and low contrast); high sensitivity ( targets in shadows); adaptivity to scenario

Secondary objectives:

C) vis-band multispectral (MS) with medium resolution (order of 1.4 m at nadir), aiming at PAN-sharpened multicolour imaging

- Field of regard: at least +- 300 Km from 420 km altitude and +-35° cross-track tilting.

- possibility of increasing the field of regard to +-420 km for +-45° cross-track tilting (but with degraded resolution)

- Earth’s coverage: nearly global

• Cross-track tilting : via spacecraft roll tilting at maximum speed of about 1°/sec .

• In track-limited tilting for non-synchronous TDI imaging: via spacecraft pitch rotation

• 3D (stereo) imaging: via pitch fore and aft tilting

• PAN ground resolution: 0.35 m at nadir;• Multispectral ground resolution : 1.4 m at nadir Instantaneous PAN and Multispectral swath: 11.5 km minimum at nadir; 14 km at 35° off-nadir Mean Revisit Interval : less than 2 weeks with one satellite and maximum off-

nadir angle of +-35°. Can be significantly reduced – to meet specific observation requirements- by

increasing the field-of-regard to: +-45°

AROSAT Mission objectives (3/5)

AROSAT Mission objectives (4/5)Imaging modes;- Spot and strips: a) full resolution/full swath : PAN only; MS only; both co-

registered; only some of the four MS colours; b) full resolution but reduced swath: as above b) binning/full swath (PAN and / or MS)

- Special modes a) Same TDI level applied to full images b) Different TDI levels applied to different portions of the

image c) strips with preprogrammed roll rate and yaw rate d) spot pairs with different pitch angles ( for 3D stereo)

• Operational duty: limited by the extent of the cooperating ground data station network .

• N° of data receiving stations: multiple, by inter-Agencies agreement; at least two located in Italy

• Two types of Data Receive Stations : fixed and transportable

• Download data rate: two simultaneous channels at 150 Mbps /QPSK modulated and one channel at 300 Mbps /8PSK modulated

• These channels can be available independently from each-other for a total maximum downlink datarate of up to 600 Mbps.

• Useful average per-pass downloading time per ground station: order of 360 sec. On board memory: about 400 Gbit . Number of std. square images for memory saturation: 60 MS and 30 PAN

AROSAT Mission objectives (5/5)

AROSAT Orbit Design (1/4)• The orbit plane forms a 60° angle with the

noon-midnight orbit (or 30° with the down-dusk orbit) which is very favourable from many viewpoints:

– It exploits the shadows cast by elevated objects in the first hours of the morning and the last hours of the afternoon as a 3D enhancement factor

– Even when not directly sunlit the satellite flies over zones of the Earth in twilight, making it possible to exploit the high sensitivity of the TDI – based camera detector arrays

– It sensibly increases the percentage of the orbit period when the spacecraft is directly hit by the Sun;

– One side of the spacecraft is always in shadow and the opposite side is always sunlit, this considerably simplifying the thermal conditioning ot the Units

One advantage of the chosen orbit plane laying is the greater percentage of time during which the spacecraft is sunlit: 77.5 % against a 62% for the classical 10am orbit and a 60% for the noon-midnight orbit.

This allows using fixed, in orbit deployed, solar panels, minimizing also the drag

AROSAT Orbit Design (2/4)

Exploiting the long shadows cast by elevated objects during the early hours of the day to support 3D

estimates

AROSAT Orbit Design (3/4)

• The behaviour of the Sun-vector to orbit plane normal is not much different from the angles characterizing the pure dawn-dusk orbit.

• We can count on a yearly average of around 62 %. This along with a 77.5% of the orbit period during which the spacecraft is sunlit, gives a mean solar area illumination efficiency of 0.48: and without using yokes .

Sun vector to porbit plane normal for 60° orbit

Sun vector to orbit plane normal for a dawn-dusk orbit

• The orbit is an heliosynchronous one with an altitude around 420 km, and a repeat cycle of 201 orbits in 13 days.

• The low orbit altitude was chosen to improve the ground resolution with existing space-qualified telescopes which offer a sub-meter resolution from spacecraft flying in the 600 –700 km altitude range. By lowering the spacecraft altitude to the 400+- 20 km range, one achieves a further 35% ground resolution improvement.

• Problems due to athmospheric drag, as well as those due to the integrated

effect of the Sun pressure, are coped with national, space-qualified, electric (Hall-effect) propulsion thrusters which are characterized by a high Isp and low thrust level ideal to cope with small external disturbing forces .

• The gain in resolution is paid with a lenghtening of the mean revisit intervals, which is however highly latitude-dependent

AROSAT Orbit Design (4/4)

AROSAT System Coverage (1/3)

Worldwide data

Revisit intervals

AROSAT System Coverage (2/3)

Area of primary interest

Revisit intervals

Telescope main features (1/5) • Unifocal anastigmatic three mirrors telescope (ATMT)

• Aperture: 0.7 m; (>> Ground Resolution of 0.35m (@ 420 km, at nadir)

• Swath (PAN & MS) 11.5 km (@ 420 km, and at nadir)

• PAN: 30,000 pixels (6 x 5000 pixel each)

• MS: 7500 pixels (4 x 7500 pixel eaxh)

• Max. Line Rate:15000 lps (synchronous and asynchronous )

• Bi-Directional Scanning

• Binning (in case of limiting data volume)

• Data volume > 5.6 Gbit/sec (PAN )

• Duty Cycle Mission dependent, can go up to 28 min. or more per orbit.

• Dimensions: 1850 mm (length) by 950 mm ( diameter)

• Mass: 125 kg; DC power: 250 W (both PAN and MS operating)

Telescope features: Imaging Modes (2/5) a) spot : square images of size 11.5 x 11.5 km

( 132 km^2)• b) strips : in the spacecraft scan direction, full

resolution or binning, with lenght up 300 km or more, depending on simultaneity with MS imaging

• c) special mode: imaging with two channel simultaneously (one channel with high TDI (dark objects), the 2nd with low TDI (bright objects)

Low TDI level

High TDI level

Telescope main features (3/5)

Light path, shown from entrance to detector

Metering structure

Primary mirror

Main body flexures

Second folding mirror and Focusing mechanism

Tertiary mirror and bezel

Secondary mirror

Pan detector assy, shown w/o holding structure

MS detector assy, shown w/o holding structure

First folding mirror

Telescope main features (4/5)

ATMT Scale: A.N

XZ

Primary Mirror (PM)

Secondary Mirror (SM) First FoldingMirror (FM1)

Tertiary Mirror (TM)

Second Folding Mirror (FM2)

ImagePlane

SecondaryImage

Exit Pupil

ATMT Scale: A.N

Optical design

Telescope Block Diagram (5/5)

PSUOM

Thermal Control

CEU - PAN

CEU - MS

Therm Cont

Therm Cont

PBU

DPSICU PDU

S/C

Launch vehicle • Compatibility with the Falcon-E launch vehicle is important to

reduce the launch cost. However compatibility with other launch vehicles will be taken into account as a back-up

The Falcon is a partially reusable launch system designed and manufactured by SpaceX. The two-stage-to-orbit rocket uses LOX/RP-1 for both stages, the first powered by a single Merlin engine and the second powered by a single Kestrel engine. It was designed by SpaceX from the ground up and is the first successful fully liquid-propelled orbital launch vehicle developed with private

funding.

• Structure: has an an hexagonal cross-section. Supports the telescope assembly through stress-relieving devices. The interface with the L.V. is on the telescope opening side.

• Thermal control: passive for most equipments, active for critical components; thermal superinsulation for the telescope

• Propulsion module: on the spacecraft rear-face includes propellant tanks, thrusters, valves. Based on electric thrusters throughout for both drag and Sun pressure compensation and for fine orbital trimming at b.o.l.

• The use of electric propulsion will be based on both the HT-100 (8 to 10 mN thrusters) and the HR-100 ( 125 mN thruster) by ALTA, and a cylindrical tank containing up to 30 lt. of Xenon.

Spacecraft Design concepts (1/11)

Spacecraft Configuration (2/11)

Spacecraft Configuration (3/11)

Spacecraft Configuration (4/11)

• Solar array: four panels folded onto the hexagonal cross-section body and in-orbit deployed , providing some 900 W useful mean orbit power;

• Energy storage: Li_ion batteries , 700 Wh nameplate capacity • Power distribution: DET concept, unregulated primary bus distribution• secondary voltages generation is responsibility of each subsystem;

• OBDH: decentralized approach. Uses a CAN bus as spacecraft bus • Supervises the operations of the payload electronics and of all subsystems; • Detailed RTU implementation : responsibility of each subsystem

• Orbit control: uses exclusively electric propulsion.• Two different thruster types: HT-100 xenon thruster (10 mN ) for drag and Sun

pressure compensation; and HR-100 thermal xenon thruster (125 mN) for orbit injection correction and fine orbit trimming

Spacecraft Design concepts (5/11)

Spacecraft Design concepts (5/11)• Attitude Control: uses four RW and three magnetic torquers as

actuators;• Attitude measurement and estimation: - uses two very accurate star sensors - two semi- coarse Sun sensors for reference and to support safe

mode; - a triaxial magnetometer as coarse attitude sensor in eclipse and

also to drive the torquerods during reaction wheels unloading - a three-axis gyro to support fine tracking in normal mode and also

to measure spacecraft rotation rates in eclispse and other non-nominal modes;

- a GPS receiver provides orbit position data to support attitude determination by other sensors, and timing to all spacecraft functions;

• TT&C: at S_band, uses multiple antennas arranged on the spacecraft to provide a nearly 4 coverage to cope with loss of attitude;

• Mass memory: modular, based on solid state, with simultaneous write/read capabilities at different data rates. Built-in logical and physical redundancy. Provides multiple simultaneous in/out accesses;

• High speed transmission system at X_band: uses two channels and two directive, repointable, commandable, antennas cross-connected to the channel amplifiers.

• Modulation is 4PSK/150 Mbps each (two channels in the EESS band) and 8PSK /300 Mbps (one channels in the EESS band).

The X_band redundant transmitter is based on a 6 W S.S. amplifiers. Uses two, 30 x 30 cm square, slotted, antennas equipped with 2 d.o.f. pointing mechanisms,

Spacecraft Design concepts (6/11)

Spacecraft mass budget

Optical payload: 120 kg Structure and mechanisms: 80 “ Thermal control: 10 “

Power and DC harness: 55 “ AOC: 40 “

Data Handling: 15 “ Propulsion ,dry: 15 “

TT&C, incl. Antennas: 10 “

Data stor.and Transm.: 55 “

Propellant ( Xenon): 25 “ Sub-total dry 400 “ Total w/o margin 425 “

Total w. 10% margin 460 “