MISSION CONFIGURATION

1. Orbit and Launcher

The ideal ASTROGAM orbit is an equatorial LEO of altitude 550-600 km. Particle background properties are ideal for this orbit as already determined by the AGILE mission (which has a LEO orbit of altitude 520-550 km and 2.5 degree inclination with respect with the equator). An equatorial orbit (required to have an inclination of < 2.5 deg, and eccentricity e < 0.01) will make use of the ESA ground station at Kourou as well as the possible use of the ASI Malindi station in Kenya. The foreseen launcher for ASTROGAM is VEGA.

2. Spacecraft

The ASTROGAM system is composed by a satellite in a Low Earth Orbit and a ground segment that includes the stations of Kourou (French Guiana) and Malindi (Kenya) (in charge of performing the spacecraft control, monitoring, and the acquisition of scientific data). The ASTROGAM spacecraft has the purpose to observe the sky according to a predefined pointing plan uploaded from ground. Different pointing profiles can be selected in order to observe selected sky regions or to perform a scanning that, thanks to the wide P/L field of view, can cover almost the whole sky at each orbit.

The spacecraft platform is made of a structure that mechanically supports the ASTROGAM instrument and hosts internally the payload electronic units and all the platform subsystems. Deployable and steerable solar panels are required to minimize the satellite envelope during launch and to allow the maximum flexibility for the payload observation profile. Figure 1 shows the spacecraft configuration with deployed solar arrays. In order to guarantee the tolerance to a single failure and to increase the reliability, the platform is fully redundant as well as the PDHU and PSU. The P/L detector modular design ensures that a single failure of one element will cause only a small reduction of the overall performances.

The total satellite dry mass (with system margin) including payload and platform is 860 kg. The required peak total power (with system margin) is about 1000 W. The average telemetry budget is 1202 kbps.

A precise timing of the payload data (1 μs at 3σ) is required to perform a proper on ground data processing able to guarantee the scientific performances of the mission. As already implemented in current missions, the required timing performance can be obtained by a GPS unit directly connected with the PDHU in order to allow a fine synchronization with the time reference.

Figure 1

Figure 1 The ASTROGAM satellite configuration with deployed solar panels and radiators.

The spacecraft is able to provide the following attitude pointings to support the payload observation requirements:

  • nearly inertial pointing (with the possibility to slowly rotate around the payload boresight) to observe continuously a selected area of the sky;
  • zenith pointing to perform at each orbit a scan of the sky;
  • fast P/L repointing during eclipse periods to avoid the presence of the Earth in the payload FoV (allowing 2 pointings per orbit).

The required pointing accuracy ( 1 deg), stability (0.01/s), and attitude knowledge of 1 arcmin (to be reached after ground processing) can be obtained using standard class sensors and actuators. The 3-axis stabilized attitude control is achieved mainly using a set of 4 reaction wheels used in zero momentum mode ensuring the possibility to perform fast repointing manoeuvres. Magnetic torquers are provided to perform wheels desaturation and to support a safe attitude pointing based on a basic subset of ACS items.

Attitude reconstruction is based on star trackers outputs; no gyros are required. Three star trackers are provided in order to guarantee the single point failure tolerance and to ensure the availability of at least one star tracker almost in any attitude condition. In addition to the star trackers, magnetometers and coarse star sensors are available.

Figure 2

Figure 2 (Left panel:) the ASTROGAM satellite in stowed configuration on top of the adapter and within the VEGA launcher. (Right panel:) the satellite with the solar arrays and deployable radiators in launch configuration.

A deployable and steerable solar array, composed by two wings, is provided to guarantee sufficient power generation in all the expected payload pointing scenarios. For the solar array, the power required at EoL is about 1900 W including a margin of 20%. The average orbital contact time with the ground stations of Kourou and Malindi is about 20 minutes for an orbit altitude of 550 km. In order to download all data, a minimum downlink data rate of about 5.7 Mbps is required. These values are compatible with the resources provided by the spacecraft platform of 10 Mbps of downlink data-rate, and 1 Gbyte of mass memory.