arosat mission and spacecraft configuration_rev3

AROSAT Mission AROSAT Mission and Spacecraft and Spacecraft ConfigurationConfiguration

Phase 0 Conceptual StudyGiorgio Perrotta, SpaceSys

Nov./Dec. 2009

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AROSAT Mission objectives Primary objectives: A) achieve Very High Resolution (<0.5m ) imaging of

artifacts ,infrastucture, small fixed or mobile objects B) Priority image features: high dyamic range (targets

with high and low contrast) in black and white ; high sensitivity;

Secondary objectives: C) vis-band multispectral with medium resolution,

aiming at PAN-sharpened multicolour imaging D) SWIR imaging at a resolution compatible with a

common telescope Field of regard: at least +- 420 Km , via spacecraft

cross-track tilting. Cross-track repointing rate : better than 0.5°/sec

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AROSAT Mission objectives

No use of tiltable mirrors for imaging purposes Instantaneous swath: option a): 6 km minimum from 420 km altitude option b) : up to 18 km (goal) “ “ “ Imaging modes: square , rectangles or long strips Earth’s coverage: global Revisit interval of any site within the Earth’s coverage

: <=1 day, with 3 satellites in the System Multispectral ground resolution (optional) : tbd

(nominally 2 x 2 m.) Typical swath width: order of 15 km;

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AROSAT Mission objectives

Operational duty: limited by the ground data station datarate capability and on-board memory;

Download data rate: 300 Mbps per channel (two simultaneous baselined). Useful per-pass dowloading time: order of 360 sec.

on board memory: 2 x 300 Gb (hot redundant) N° of images for memory saturation: option a, (10 x 10 kpixel): 150 odg option b, (30 x 30 kpixel): 60 “ N° of data receiving stations: multiple, by inter-

Agencies agreement; at least two located in Italy Data receive stations: can be fixed and transportable

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AROSAT System Design

The system-level revisit interval requirement implies three spacecraft injected in the same SSO orbit at an altitude of 412 km, and 120° staggered in anomaly.

The orbit plane forms a 60° angle with the noon-midnight orbit ( or 30° w.r.t. the down-dusk orbit) which is favourable from many viewpoints.

First, it exploit the shadows cast by elevated objects in the first hours of the morning and the last hours of the afternoon. Second, it increases the portion of the orbit when the spacecraft is sunlit. Third, even when not sunlit it overflies zones of the Earth experiencing twilight making it possible to take pictures using TDA-based linear arrays as high sensitivity camera sensors..

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AROSAT Orbit Design

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

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AROSAT Orbit Design

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

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AROSAT Orbit Design

The orbit is an heliosynchronous one with an altitude of 412 km, performing 31 orbits in two days.

The low orbit altitude was chosen to improve the ground resolution with whichever existing high-resolution space-qualified telescope for which a sub-meter resolution is already offered from spacecraft flying in the 600 –700 km altitude range.

Problems due to drag will be coped recurring to national space-qualified electric (Hall-effect) thrusters developed in Italy;

Putting and mantaining equispaced all three spacecraft in the same orbital plane simplifies considerably the launch and orbital control of the small constellation.

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AROSAT System Coverage

The 3-hours FOV projection on ground from the three spacecradt show how they achieve a contiguity of their fields of regard.

Accordingly a daily revist of any site on Earth is feasible using the spacecraft roll rotation to position the optical instrument swath inside the field of regard.

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Ground tracks of the three satellites over 24 hours

Ground tracks of the three spacecraft over 12 hours

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Projection of the fields of regard of the three spacecraft for a time interval of 12 hours

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Telescope Options Two alternatives at present : A) a VHR panchro only telescope: meets most of the

priority requirements. Swath ~ 6 km.from 412 km altitude Dimensional approx. envelope : 1520 x 790 diam. Mass: 50 kg o.o.m.; DC power drain: 50 W o.o.m.

B) a VHR panchro + multispectral (no SWIR) telescope: meets priority 1 and 2nd rqmts. Swath ~15-18 km from 412 km altitude. Dimensional envelope: 1850 x 950 mm. Mass : 125 kg; DC power drain: 200 W o.o.m.

Strong impact on data storage and transmission of option B) vs option A)

Spacecraft conceptual design for options A and B: similar but not identical , requires ad hoc ‘personalizations’.

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Spacecraft Design concepts (1) Compatibility with the national VEGA launch vehicle

is important to the light of ASI investments in the programme. Finding a solution enabling a single launch for all three spacecraft will be a cost-attractive solution.

The VEGA inner shrouds diameter is 2380 mm. It can accommodate three spacecraft each characterized by a circular envelope with a diameter of 1100 mm, or four satellites with an external circular envelope of 980 mm diameter

The telescope diameters above listed allow launching together three spacecraft version A) or B) with a single VEGA launcher

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Spacecraft Design concepts (2)

Structure: a truss structure with an hexagonal cross-section. It 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 and propulsion assembly (tanks)

Propulsion module: on the spacecraft rear-face, it includes propellant tanks, thrusters, valves. Chemical or electric alternatives with different impact on mass, performance;

Solar array: six panels folded onto the hexagonal cross-section body and in-orbit deployed. The S.A. is not Sun-tracking ;

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The spacecraft options- A with an estimated drymass of 290 kg, and B with an estimated drymass of 350 kg - have nearly the same ballistic coefficient (about 7 and 7.57 10^-3).The low effective area is due to the fact that the solar panels are parallel to the orbit plane, minimizing the drag;

Assuming a launch in 2012 we face a period of decreasing Sun activity. Nevertheless, to counter drag an average velocity increment of 58 m/sec/ year is needed.

Using chemical (hydrazine) or electrical (xenon) propellant its mass, for 7 years operation, is:

chemical M : 61 (sat. A) ; 75 (sat. B) electrical M : 12 “ ; 16.5 “

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The chemical propulsion has a negative impact on the spacecraft configuration. Indeed a tank with 50 to 60 cm diameter would be needed; or else from two to four smaller spherical tanks would be required still with a spacecraft height increase of 20 to 40 cm.

The use of electric propulsion looks more favourable based on 10 mN thrusters by ALTA (Italy) and a cylindrical tank (approximate size 40 x 10 cm, superinsulation included) containing up to 30 lt. of Xenon.

The operativity of the thrusters, at present under qualification, for a continuous service of greater than 40000 hours, remains to be assessed.

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Energy storage and power management: Li_ion batteries, semi-distributed power regulation and secondary voltages generation;

OBDH: decentralized approach, controls the essential spacecraft functions and supervises the operation of the payload electronics and its interfaces with other subsystems;

Orbit control:via the propulsion subsystem under control by the OBDH also using GPS receiver data;

Attitude Control: uses RW and magnetic coils as actuators.

Attitude measurement and estimation: based on star sensors, Earth, coarse Sun, magnetometers sensors, a gyropack and the GPS receiver data;

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For attitude determination and control , two options are possible by proper choice of the star sensors:

- a 35 m GSD which may be sufficient for area contours georeferencing and medium accuracy location of human and natural artifacts ;

- a 2.2 m GSD which can be used for precision location and targeting of human artifacts for defense purposes or for special civil protection tasks

these options impact differently the other Subsystems and costs

Precise objects location determination is also software intensive too;

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Spacecraft Design concepts (8) The power budget shows an average demand for bus

and payload, between 450 and 700 W. The Solar array consists of six rectangular panels

in-orbit unfolded by means of simple hinges. Panels size is: version A: 1.6 x 0.45 m; version B: 2.0x 0.5 m. Low-grade GaAs cells ( of 20% ) are used throughout to reduce costs.

the generated power, averaged over the orbit, is: version A) : ((6x 0.72) x 0.77 x 0.86 x 0.20 x 0.85) x

1366= 664 W version B): ((6x 1.0) x 0.77 x 0.86 x 0.2 x 0.85) x 1366

= 922 W

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

High speed transmission system at X_band: uses two channels and two directive, repointable, antennas cross-connected to the channel amplifiers. Modulation is 4PSK (one channel in the EESS band) or 8PSK (two channels in the EESS band), depending on data downloading operational requirements

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;

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The X_band high data rate transmission system is an outstanding spacecraft subsystem. The block diagram is shown below, with two independent transmission chains, each carrying a wideband (300 Mbps) moulated carrier centered on different frequencies and linked to independently repointable directive antennas. The latter can transmit the same, or different, data towards the same, or different, destination stations within the istantaneously available but time-variable access area.

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Spacecraft Design concepts : the HDRT (12)

6 W S.S. amplifiers, 30 cm square, slotted antennas equipped with 2 d.o.f. pointing mechanisms. Fixed or transportable ground antenna terminals, of 2.5 m diameter, assure reception of one or two 300 Mbps channels within the spacecraft instantaneous coverage.

Higher per-channel data rates can be provided (growth capability) using a combination of: 8PSK, larger ground antenna terminals (say: 5 m diam.) , efficient coding systems, greater spacecraft TX power.

The spacecraft-to-ground X_band link could be complemented by a low datarate X_band uplink channel for secure command and auxiliary service communication functions (not in the baseline )

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Spacecraft views version A

The telescope rests on two plates and is supported by a truss-type structure. The propulsion module is supported by a separate platform also anchored to the truss-structure

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Telescope anchored to a truss type structure made of CFRP tubes and joints. The electric propulsion system is supported by a thick honeycomb/CFRP sandwich

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A set of honeycomb panels support the electronic boxes and close the truss-type structure this resulting in a stiff composite.

The panels can be individually assembled and thermally controlled depending on the hosted equipments

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The composite carries 6 solar panels which are folded onto the six faces of the hexagonal cross section satellite body, then released and deployed once in orbit

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When unfolded the solar array forms two wings that are kept parallel to the orbit plane, this minimizing the drag while providing a fairly good illumination of the array for the 77% of the orbit period. All six spacecraft panels see the open space and can dissipate heat where it is more convenient.

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This spacecraft top view shows the symmetry of the solar wings and their geometrical relations to the spacecraft body. The Hexagonal shape avoids completely any shadowing for the chosen orbit plane laying, since the Sun vector makes an angle of 30° w.r.t. the normal to the Solar Array

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Spacecraft version A: mass and power budget

Optical payload: 50 kg opt. Pay. Electronics 50 W Structure and mechanisms: 50 “ battery charging 100 “

Thermal control: 10 “ T.C. 50 “ Power and DC harness: 60 “ Pow. 40 “ AOC: 40 “ orbital average 40 “ Data Handling: 15 “ D.H. 20 “ Propulsion ,dry: 10 “ orbital average 100 “

TT&C, incl. Antennas: 10 “ orbital average 10 “ Data stor.and Transm.: 45 “ orbital average 150 “

sub-total dry 290 “ subtotal 560

Propellant ( Xenon): 15 “

total with 10% margin 335 “ total w. 10% margin 616

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Spacecraft version B

Spacecraft version B is larger than version A since it carries a larger and more powerful Telescope. Main differences, w.r.t. the telescope of version A, are:

- a larger aperture >> better ground resolution; - a larger PAN swath (3 times that of vers. A) - besides a panchro channel it also implements a

multispectral system with four frequency bands in the visible

These enhanced features impact spacecraft size, mass and power, mass memory and D.H., operational flexibility, tthe attitude control, and the tranmission system.

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Spacecraft views version B

This is a sketchy view of the telescope which is connected to a truss-structure by means of honeycomb panels

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This view shows the truss-type skeleton supporting internally the Telescope.

Honeycomb panels strenghten the truss assembly.

The Solar Array panel are wrapped around the hexagonal body and released in orbit by means of preloaded spring actuated hinges.

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This view shows the main spacecraft skeleton with the side panels, all removable to install the electronic Units.

The panels modularity allows to indivudually optimize their thermal balance .

The panels provide ample room for allocating S_band TTC antennas and the GPS antennas

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The top floor carries the Electric Propulsion system (one Xenon tank, and two HT-100 thrusters), and two star sensors tightly connected to the telescope own structure

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Two mechanically steerable antennas (30 x 30 cm) are fitted on the Earth facing panel for transmitting X_band high speed data to two different ground terminals within the satellite field of view

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Spacecraft version B mass and power budget

Optical payload: 120 kg Opt. Pay. Electronics 200 W

Structure and mechanisms: 60 “ Battery charging 120 “

Thermal control: 10 “ T.C. 50 “ Power and DC harness: 70 “ Power plant 50 “ AOC: 40 “ AOC orb. Aver. 40 “ Data Handling: 15 “ D.H. 20 “ Propulsion ,dry: 10 “ El. Prop. aver. 150 “

TT&C, incl. Antennas: 10 “ TTC orb. aver. 10 “ Data stor.and Transm.: 55 “ DST orb.aver. 180 “ sub-total dry 390 “ subtotal 820 “Propellant ( Xenon): 20 “ margin 10% 80 “

total with 10% margin 450 “ total with margin 900 “

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Performance of a system with one satellite only Parametric, though preliminary, analyses were

performed in case the System would be based on one spacecraft only.

The coverage is still global but the mean revist interval - with a field of regard delimited by a roll axis tilting of +-45° - increases from 1 (case of three satellites) to 3 days.

Small performance variations can be achieved by changing the orbit altitude within a +- 35 km w.r.t. the nominal value of 412 km,

There are a few critical zones for which a 3-days revisit interval can become marginal. A slight increase of the field of regard plus small altitude changes may be considered to improve the situation, if needed.

arosat mission and spacecraft configuration_rev3

Documents

spacecraft views version

spacecraft version b

power plant

amultispectral system

thetranmission system

aoc orb

ttc orb

theside panels