“Immediate” return version
In this case the MCT stays between 30 to 60 days, at most, on the surface of Mars and undertakes its return flight along an “opposition” pattern trajectory, allowing it to reach Earth before the next launch window, thus maximizing its productivity. This trajectory is actually an ellipse with a perihelion below Earth orbit (close to that of Venus), the achievement of which requires an additional ΔV compared to the return along a Hohmann trajectory (which we considered so far), of about 1.8 km/s, bringing the total impulse to deliver up to 8.32 km/s (Mars at aphelion). For a 380 s Isp (Raptor adapted to vacuum) this induces a mass ratio (between mass at takeoff and mass injected into transfer) of 9.3, that is to say a mass of propellant representing 89% of the total mass at takeoff. Such an important value is at the limit of what can be achieved by the shuttle; it leads to a mass of propellant largely dimensioning for the size of the tanks and the thrust of the engines. Spherical engines would not fit into the 20m wide fuselage; it is required to get back to cylindrical tanks with common intertank bulkhead, split, complicating the drawing conditions.
This return also sets more constraints onto the travelers because it is longer (11 months) and leads closer to the sun, one of the sources of ionizing radiations. However since, in a colonization scheme, the number of people on board for a return flight is by construction assumed to be small (typically a dozen), it will be possible to protect them effectively by concentrating around their smaller living quarter the anti-radiation protection panels designed as a screen for the 100 passengers of the outward leg.
Finally, as already pointed out just before § 2.1.1, this trajectory imposes a very serious challenge in terms of design of the TP, due to a very high return speed.
|Length (m) (without base hood)||44|
|Payload mass, outward leg (T)||82|
|Payload mass, return leg (T)||5|
|Structural coefficient, Tanks||0,040|
|Prop. margin mass coefficient||1,030|
|Prop. 3rd stage LEO complement (T)||616,5|
|ΔV outward leg, Hohmann Mars at perihelion||3.46|
|ΔV return leg, opposition (immediate return)||8.32|
|ΔV maneuvres Mars orbit||0.15|
|ΔV EDL braking & hovering||0.60|
|ΔV maneuvres Return to Earth||0.30|
|MASS ITEMS (T)|
|6 of 120 T||7.0|
|Structures & Thermal Protection||60.0|
|STATUS of TOTAL MASS (T)|
|On LEO parking||691.5|
|During Earth-Mars transfer||265.5|
|At EDL Entry||247.6|
|At taking off from Mars||1231.0|
|PROPELLANT MASS (T)|
|Injection transfer Earth-Mars||426.0|
|Maneuvers Mars orbit||17.9|
|Maneuvers Return Earth||11.0|
|Tank Capacity (as a function of the return flight)||1103.3|
The payload mass must be curtailed again, down to 82 T, if we do not intend to increase the thrust of the launcher at takeoff (number of engines). Despite this, the payload to be placed on LEO is larger than in the previous case, because the tank capacity, this time determined by the return flight, is significantly higher (1,103 tons against 865), which results in a more penalizing dry mass of the vehicle (122.7 tons versus 110.6). On the other hand, this increased capacity allows to allocate a larger quantity of propellant at the end of the insertion into LEO by the shuttle itself (617 tons instead of 400), which explains why the launcher mass on LEO reaches 691, 5 tons.