VHR_preliminary-system_study_issue2

Preliminary AROSAT system study

RHI, 16 –17 Nov. 2009

By: G.Perrotta / SpaceSys

Preliminary requirements (1)

• Coverage: temperate latitude belt up to +- 50° (tbc). The coverage of the polar caps is: optional

• Fully operational system: two or three satellite system, all operating. Graceful degradation with just two spacecraft operational

• Orbital planes: tbd, under evaluation;• Pancromatic camera with better than 0.5 m grd resol. at nadir• Images: square, around 6 x 6 km (goal: 12 x 12 km) ;• continuous strips 6 km wide (goal: 12 km) x 70 km long• Optical beam repointing: up to +- 35° cross-track • Optional: up to +- 30° along track fast repointing ( for 3D

imaging)• Operational duty: nominal average of 0.2% on a per-orbit basis,

with a maximum of 1 % on one orbit per day;

Preliminary Requirements (2)

• On board memory : compatible with 50% (tbc) of the data acquired daily;

• Gathered data download data rate: < 150 Mbps, 1 or 2 channels

• On-board data download antennas: 1 or 2,directional, repointable, providing simultaneous or independent operation;

• Ground receive data stations: multiple, provided with antennas with diameter in the 4 to 5 m range;

• Full system: compatible with single VEGA launch and other vehicles capable of multisatellite launches;

• Single spacecraft (demostration flight): launch vehicle and site is: tbd

• Replacement or substitution satellite(s): under study

Impact of Requirements • We face contrasting requirements:

a) To point the optical telescope beam at nadir while allowing cross-track or along-track repointing, on demand;

b) To extract from the Sun the maximum energy possible minimizing the costs;

c) To minimize the spacecraft and appendages cross-section area to minimize drag;

d) To go down as fas as possible in altitude to increase ground resolution while minimizing the telescope size and mass;

• Besides, the cost reduction pressure by Customers push towards

avoiding unnecessary sophistications

Alternate spacecraft configurations (1)

Nadir pointing

Sun vector

Pitch axis

This conventional approach deploys two solar wings which are kept Sun-pointed via slip-rings throughout the orbit portion wherein the spacecraft is sunlit.The telescope axis is , normally, nadir pointed.

This configuration presents a large cross-section to the air flow and is not the ‘optimum’ for low-altitude flying spacecraft. A positive factor is that the solar panels can be kept sun-pointed thoughout the ‘active’ orbit portion, therefore the solar cells are efficiently used. To achieve this the solar panels must be equipped with a BAPTA and drive system, which are critical and costly items.


Pitch axis

Sun vector

Optical beam direction

Specular flat surface

In this configuration a 45° flat mirror redirects the optical beam towards nadir, Te cross section area-affecting the drag force- in minimized also because the solar wings are edge-seen, with a positive impact on mass and propulsion system

The solar arrays are folded onto the spacecraft body and in-orbit deployed. The solar energy capture varies throughout the orbit thus deceasing the mean energy collection.However the lower efficiency is counterbalanced by a much simpler solar plant, needing only simple hinges to deploy the solar wings.


Sun vector

Roll axis = velocity vector

45° inclined flat reflector

Nadir-looking optical beam

This is the previous configuration rotated by 90° . The side solar panels are provided with motorized hinges so that during the near polar portions of the orbit one can get more power from the Sun rays than with the previous configuration.

The advantages, to be better quantized depending on specific mission requirements, are counterbalanced by a greater cross sectional area impacting the drag, which results to be of the same order of magnitude ot the first configuration.


• Assumptions: spacecraft body lenght x width: 1.5 x 0.6 m• Solar panels: 1.5 x 0.6 m eachComputed mean solar panel effective area (over the sunlit orbit

portion) : Conf. #1 = 2.7 m^2Conf. #2 = 1.9 m^2Conf #3 = 2.15 m^2

Computed equiv. Area for drag during the sunlit (*) orbit portion: Conf. #1 = 2.7 m^2Conf.#2 = 0.42 m^2Conf. #3= 1.55 m^2

(*) in eclipse all config. are modified to minimize air drag

Preferred solutionAe for drag Solar array

effectivenessRisk and maturity

Config. #1 High Good Reference

Config. #2 Very Low Medium Equivalent to reference

Config. #3 Medium to High

Slightly better than medium

A bit worse than config. #2

Configuration # 2 can be advantageously pursued from a drag (and propulsion) viewpoint , Solar Array type ( simple mechanisms) and related cost savings. The flat specular mirror will have to be designed and built with utmost care, however.

Basic formulas for optical instrument • IFOV = Rs/H ( Rs= ground resolution; H= satellite altitude)

• F = ps/IFOV ( F= focal lenght; ps= pixel size)

• Do = 1.22 *F* /ps ( Do= pupil diameter; = wavelenght)

• Telescope size is mainly affected by Do and F.

The equations can be rewritten as:

• Do = 1.22* * H/Rs : a low H reduces the pupil size

• F= ps* H/Rs : the smaller the pixel size the better

• Most existing and planned satellite fly high to avoid counteracting the drag using the heavy chemical propulsion. But doing so increases F and Do with attendant significant cost and mass increases.

• Lowering the orbit altitude restores manageable values for telescope and spacecraft mass, power and cost, but this cannot be cheaply achieved with chemical propulsion (high mass impact) so one has to resort to electric propulsion

Impact of spacecraft altitude on camera parameters

0

2

4

6

8

10

12

altitude, km

foca

l len

ght,

m

pupil diameter for 0.5 m grd resolution

0

0.2

0.4

0.6

0.8

1

1.2

350 400 450 500 550 600 650 700

orbit altitude , km

diam

eter

, m

The Focal lenght is valid for a specific value of the pixel size, here 7 microns

Athmospheric drag compensation

• Alternatives:

• Chemical propulsion, Isp of 210 sec

• Electric propulsion, Isp of 1000 sec. (conservative, Hall-thrusters)

• The drag was computed for 7 years starting from 2012 assuming mean values for F10.7 and Kp coefficients .

The ballistic coefficient was computed for different spacraft configurations

The V required to counteract the drag was then computed vs orbit altitude and S/C configuration.

Athmospheric drag compensation

• From published forecasts of the expected Sun activity for the next Sun cycle, which seems to be considerably more quiet than the previous one, a mean F10.7 value around 120 can be computed. This was increaded to 140 for margin

• With these assumptions the 5 years velocity increment required to counteract the drag effect for the three configurations was computed

• For the two propulsion systems the propellant mass result as follows:

chemical electrical• Conf # 1 163.6 kg 29.8 kg• Conf. # 2 18.9 “ 3.9 “• Conf. # 3 74.8 “ 14.7 “These data show that Config. #2 could be a valid candidate in all

cases

Propulsion system considerations

• The data shown above imply that Configuration #2 could be supported by a less costly chemical propulsion system, while Configuration #1 would likely have to be equipped with electric propulsion not to increase too much the launch mass and spacecraft size that could make difficult the compatibility with less expensive launchers, VEGA included;

• Nevertheless other requirements, including allowances for ‘missions of opportunity’ requiring orbit changes, or demanding orbit injection manoeuvres, and a more precise assessment of the spacecraft dimensions and mass – depending from the optical instrument configuration and features- might lead to different conclusions;

• The use of a mixed chemical and electric propulsion cannot be excluded at this stage and will be reconsideed in the following;

Spacecraft architecture (1)

The satellite #2 architecture would include a cradle supporting the telescope, a propulsion module (in yellow) carrying either the chemical or the electrical propulsion items, a frame supporting a 45° inclined flat mirror, and the solar array in three panels, two of which stowed during launch and deployed in orbit by means of motorized hinges


This is a pictorial view of the spacecraft with the solar array deployed. Heat rejection can occurr via the side panels. The back of the 45° mirror is available for carrying an ISL package for data relay to a tbd satellite system. Two additional panels could be installed between the mirror and the cradle to support high datarate X_band link with Earth, and also at S_band for TT&C and GPS reception.


• The spacecraft electronics is rather conventional with few exceptions.

• The short term pointing stability implies the elimination or reduction of microvibrations, and suitable technologies will have to be used;

• The image center should be within a 10% of the image size which is felt to be sufficient for surveillance / defense tasks in populated areas. This would imply an angular pointing indetermination of 1.5 mrad, implying the use of star sensors;

• The X_band transmission system would be based on 8PSK modulation, effective coding and the use of on-board compression systems. The approach of having two transmission chains centered on different center frequencies and radiating through two separately pointed directive antennas, provides operational flexibility, redundancy and graceful degradation;

• The use of multiple TT&C S_band and of GPS antennas for signal reception, is also envisaged to widen the accessibility coverage;

Spacecraft High Data rate transmitter The high data rate transmission system is an outstanding spacecraft subsystem. A candidate block diagram is shown below, with two independent transmission chains, each carrying a wideband moulated carrier centered on different frequencies and linked to an independently repointable directive antenna 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.

Spacecraft budgets

• Spacecraft mass and DC power budgets cannot be properly elaborated without a precise knowledge of the Camera geometrical envelope and of the key electrical interfaces. The Camera lenght has, indeed, a deep impact on the structure , solar array and other subsystems.

• However, and for reference purposes, we have estimated the mass budget for the Camera version that we propose for an advanced implementation based on a 0.5 m diameter pupil, a focal lenght around 5 m , an APS-CMOS detector matrix with 7 micron pixels. We assumed a spacecraft flying around 400 km and provided with chemical propulsion system (worst case). The mass depends on the redundancy implementation but we are well below the target of 400 kg at launch.

• Once the Camera will have been chosen, we plan to reconsider both the configuration choice ( either the #2 or the more conventionall #1) and the overall mass budget.

System coverage (1)

The coverage, intended as extension of the access area, has been assessed with one, two and three satellites. The access area limits was taken, respectively, 30° and 40° half-cone angle. With one satellite the access are is extremely limited and does not allow to decrease the revisit interval below 3 or 4 days. The situation improves with 2 satellites and can be considered satisfactory with three spacecraft .

System coverage (2)These graphs show the access area improvement feasible using three satellite spaced by 120° in true anomaly,same orbital plane. The satellite must rotate in roll by, respectivly, +- 30° ( with a resolution loss of 22% in the cross –track direction) or +- 40° ( with a resolution loss , cross-track, of 30% ) . This system configuration allows a revisit interval of 1 day for almost all points of the Earth, except small areas close to the poles.

VEGA compatibility •The system coverage can be achieved incrementally, launching satellites one-at-the-time.

•However if achieving a very short revisit interval is a priority issue, then multiple simultaneous launches are mandatory.

•In case of a three-satellite system they could be accommodated inside the VEGA shroud and launched together: but the satellites cross sections should be kept (see the side picture) within a 650 x 950 mm envelope.

The VEGA inner shroud diameter is 2380 mm and can accommodate one satellite or even two spacecraft with cross-sections around 1.1 x 1 m.

VHR_preliminary-system_study_issue2

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