Asteroid Resource Recovery Proof - of - Concept & Sample Return Mission

The simplest‘ proof-of-concept’ asteroid mining mission is identified to be one which simply seeks to demonstrate collection of bulk sample for return to Earth for geotechnical (ie strength) and metallurgical (ie product recovery)testing. In the mineral industry world, this mission should be viewed as a ‘trial mining demonstration& metallurgical sample collection’ campaign.

This mission, as it is facilitating test work, and demonstrating key technologies, would be regarded interrestrial mining project planning terms, as a key component of a detailed Feasibility Study for planning a full-scale commercial project.

It is proposed that an adequate sample size to support such studies for future commercial missions, and to ‘feel right’ as a demonstration, is about 50 kg, but not less than (eg) 10 kg.

Since such a Proof-of-Concept mission is justified as a facet of a larger Feasibility Study, commercial break even is not required. Nevertheless, there is likely tobe significant earnings from sale of samples to collectors and scientific institutes, and sale of ‘video and story content’ to media organizations, and sale of advertising visibility for sponsors.

Given that the propellant for mass return has to come from Earth, we need to identify targets which have exceptionally easy accessibility (delta-v outbound, including rendezvous) and exceptionally easy return (delta-v asteroid-depart and relatively easy earth-capture or relatively low direct re-entry velocities). These are low-eccentricity and low-inclination Apollo or Aten asteroids.

There are three distinct earth-return scenarios: (a) capture into a HEEO following lunar flyby (LFB) and then gradual reduction of orbit apogee to a LEO, with rendezvous with the International Space Station(or other manned LEO space station)and then return to earth via a supplied service, or(b)re-entry from HEEO or the achieved LEO; or(c) direct earth re-entry from asteroid hyperbolic return trajectory. Re-entry would presumably be targeted eg to Woomera or into an ocean recovery target area. A recent additional possibility would be return to the future Lunar Gateway station.

Direct re-entry is less constrained, in the sense of limiting trajectory parameters such as v-in f and z-vector on arrival at Earth Sphere-of-Influence, but requires an as-yet-unknown mass allocation for aero brake. Return to LEO via LFB, then re-entry, is however more in line with the ultimate payload delivery task, and in addition demonstrates the LFB task, and is the preferred option for this PoC Mission.

Various searches for ‘good targets’ have been carried out using several approaches including interrogation of JPL Small Bodies database, and NASA Trajectory Browser, and we have successfully demonstrated to our own satisfaction that there are generally two or three ‘good’ targets each year (ie, we are not short of mission possibilities).

Target has to be a highly accessible NEA giving a short (eg under 2.5yr) return mission. ‘Highly accessible’ is here taken to be no more than about 6km/s out bound covering all impulses from LEO depart to asteroid rendezvous. But the lower the better .Return payload to be of target mass ca. 50kg (but 10 kg still acceptable). The target also, critically, must provide for low-dv returns.

An adequate dv budget com ing from review of these missions, is as follows:

The outbound dv budget(being for LEO-depart +injection into heliocentric transfer + asteroid arrival)needs to be around 4 or 5 km/s, for ballistic (impulsive) departure from LEO. If departure from LEO is to be via low-thrust non-impulsive deep space ion / Hall thruster, then the outbound dv budget needs to be roughly double the impulsive demand, ie about 9 km/s. This strongly suggests departure via direct injection using chemical thruster, via perigee Oberth burn, most likely from a GTO or super GTO drop-off. Note that the arrival dv can readily, and best, be supplied by the deep-space electric(eg Hall Effect) thruster.

The return dv (being asteroid-depart + any pre-Earth-Sphere of Influence-arrival thrusting+ any post-lunar-flyby manoeuvres) must be below about 3.5km/s

The biggest dv demand is in fact, killing pre-Earth-arrival hyperbolic velocity (if we want capture into orbit, rather than design for direct re-entry to Earth’s surface).

The ability to ‘close’ with a mass budget for return of ca 50 kg depends on total dv reqt and MOSTIMPORTANTLY, on the propulsion system chosen. Given that all propellant is earth-origin, highest specific impulse Isp is desirable, for maximizing impulse from limited propellant mass. An Isp of the propulsion system above 2000 seconds is appropriate. This points to use of ion or Hall thrusters.

Scope / Mission Outline ( to be tested ):

Either: Launch viaa dedicated small launcher(eg) ISRO SSLV, Rocket labs Electron, or similar, being representative of the new low cost rapid prep launchers becoming available, to provide a launch into the max energy orbit achievable (nominal for the SSLV is max 500 kg into 500 km circular).Ifearth-departureis from LEO then there will be additional demand on the deep-space propulsion system.

Or: contract with Spaceflight Services or True Orbit for ‘standby’ launch on whatever vehicle they are filling manifest of.

Preferable is launch into GTO or super-GTO as a piggyback on a mission with primary task of delivery of a large satellite into GEO; or directly to hyperbolic (ie heliocentric) on a dedicated launch

Or : As a totally different alternative: we have before us the Sky corp proposal, for launching from ISS carried on their reusable logistics vehicle, which uses2NASA Hall thrusters type103M.XL, return via the same space craftto ISS for retrieval and return to earth. Or a bespoke-engineered cut down equivalent craft..

Spacecraft ‘for example’ mass budget:

As an ‘existence-proof’ comparison, DSI was during 2015-16working on an asteroid ‘Explorer’, intended to return 10 kg from a small extremely accessible NEA to Earth, to be launched from LEO. Its dv budget was to be 5.3 km/s, using storable chem propulsion, departing from LEO. The craft wet mass was to be 315 kg, dry mass was listed to be 52 kg, propellants and pressurant was 260 kg. With in the 52 kg dry mass budget, avionics, nav, science and comms mass came in under 9 kg.

More recent info suggests this mass budget can be further significantly reduced. But you would probably use the freed-up budget to provide more dv out, so as to be able to access more targets.

We presume similar mass assumptions can be made, with the exception that there will need to be estimates for solar panel mass and power conditioning, and for the ion or Hall thrusters assembly.

Best estimates of specific power for new-generation PV panel assemblies is something like2.0 down to a most optimistic0.5kg/kW, to give (say) 4 kW then requires 8 kg

The Hayabusa-2‘mu-10’4-engine pack had mass of 66 kg. Assume the same for whatever is chosen. Or check NASA 103M.

Mass of prototype miner system to be ca 20 kg. Prototype is under development

The mass budget then looks something like total dry mass:160kg; propellant: 135 kg.

So wet mass, at LEO depart, is about 300kg.

Now, note, there’s significant slack in this WAG set of estimates….and the dv budget is significantly in excess of what we will hopefully need, depending on the specific target.