Case Study: Vega AVUM Upper Stage

The Vega AVUM will serve as a case study for a launch vehicle upper stage. As mentioned in 9.3, the final orbit and deorbiting strategy for an upper stage depends greatly on the main payload. This case study will use the orbit of the AVUM that was launched on the maiden Vega flight in 2012. The upper stage was

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left to decay from an elliptic obit with perigee at 270 km and apogee of 1250 km (Aug 2013), and inclination of 69.5°.

The mass of the upper stage is 670 kg. For this particular flight, the LARES-A&H/SS (Laser Relativity Satellite - Avionic & Harness / Support System) remained attached to the upper stage, incurring an additional 300 kg mass for a total mass of 960 kg. The drag surface area is 3 m2_{. For the purpose of finding}

the collision cross section area, it is assumed that the projected shape of the upper stage is a circle with 1.95 m diameter.

Figure 10-20 Vega launch vehicle (left) and AVUM upper stage (right). Image credit ESA No mitigation

The AVUM in the test orbit will decay naturally in 8 to 10 years. For such a decay time, the risk of a collision is around 0.018%. In this instance, the upper stage does not need a deorbiting solution to allow it to comply with mitigation guidelines, but it will be instructive to see how the collision risk can be reduced by alternative strategies.

Propulsion

The ultimate risk reduction method would be to perform a controlled re-entry using the AVUM’s re- ignitable 2.45 kN motor. The delta-v required for a controlled re-entry from the initial orbit is between 65 m/s and 350 m/s, depending on where in the orbit the re-entry burn is performed. With an Isp of 315 s, the smallest delta-v manoeuvre requires 20 kg of propellant.

Drag-sail deorbiting

With the inclusion of a 25 m2 _{drag sail the decay time for the AVUM is reduced to around 200 days. Such}

a sail sub-system will only weigh 3.5 kg. The risk of a debris generating collision is reduced to 0.002%. It is assumed that a suitable mounting configuration can be found that will result in passive attitude stability.

Inflatable sphere

An inflatable sphere with 25 m2 _{projected drag area will yield similar deorbit times as the sail scenario}

above. The time to deorbit is 240 days, with a debris-generating collision risk of 0.0012%. The deorbiting sub-system will have a mass of 33 kg.

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Tether

An electrodynamic tether with a length of 200m will again produce similar deorbit time as the drag augmentation alternatives. The tether and supporting sub-systems will weigh 18 kg, with associated debris generating collision risk of 0.0023%.

Summary: Vega AVUM

The deorbiting strategies for the Vega upper stage is summarized in Table 10-6.

The safest deorbiting strategy is to perform a controlled re-entry. But there are a number of reasons why such an option might not be used. In this test case the decaying elliptic orbit in which the stage was left is sufficient to comply with mitigation rules. It is possible that it was more desirable to use all of the available propellant for the payload insertion, and perform a shorter burn to end up in the decaying orbit, than to spend more propellant for a controlled re-entry.

Another reason for not performing a controlled re-entry is in the event of failure of the upper stage. In such a situation it will be beneficial to have a back-up deorbiting strategy.

Table 10-6 Summary of deorbit alternatives for Vega AVUM Deorbiting

strategy Detail Mass (kg) d.g. ATP (m2_.yr)

Collision probability

(%) Sub-system

requirements D.g. Op. _interfere sat.

No mitigation 8 years natural decay 0 36.61 0.018 - -

Drag sail 25m2 _{sail, 195 days 3.5 4.68 0.002 0.000 -}

Sphere 25m2 _{balloon, 240 days 33 3.01 0.001 0.000} Inflation pressure

maintained

EDT 200m tether, 260 days 18 3.75 0.002 0.006 Requires operations

input and active satellite sub-systems controlled re-

entry delta-v = m/s 20 0.00 0.000 -

Whatever the reason for choosing an alternative solution to a controlled re-entry, of all the available options the drag-sail can be implemented with lowest mass penalty. A small sail is all that is needed to deorbit the upper stage in less than a year, to result in a collision probability that is 1/10th _{of that of the}

current strategy.

10.8 Summary

The analysis methods and constraints that were derived earlier in this thesis were applied in this chapter to three case studies – two communications satellites and a launch vehicle upper stage. The analysis was extended by comparing results of a sail-based deorbiting strategy to that of conventional propulsion, electrodynamic tether and inflatable balloon.

It was found that a drag sail strategy will serve as a good back-up for the Iridium NEXT satellite. A small 3.5 kg sail can be used to deorbit the Iridium NEXT satellite in 15 years with full compliance to mitigation guidelines. But such a strategy as the primary deorbiting means will most likely be rejected by the satellite licensing authority based on a previous decision regarding propulsive deorbiting. The primary deorbiting method will continue to make use of the existing propulsion system since this has the lowest risk associated with it. The small sail is still attractive in this case as a secondary or fail-safe option.

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The Globalstar 2nd _{generation satellite can make use of an initial solar sailing orbit lowering phase}

followed by drag augmentation from the same sail, to comply with mitigation guidelines. But the required sail is fairly large – 700 m2 _{with a mass in excess of 60 kg - and other alternatives such as the}

electrodynamic tether appears to have less risk associated.

A drag-sail appears to be the best deorbiting option for the Vega AVUM, especially in instances where a controlled re-entry is not intended. A small 3.5 kg sail is sufficient to deorbit the maiden upper stage from its current elliptic orbit.

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