Chemical Rockets

Typical top-level comparison of six different architectures for transporting payloads from Earth using chemical rockets. This paper assumes LH/LOx engines with a 462 second specific impulse, and a 200 km circular reference orbit inclined 51deg.. The six architectures examined are:

1. Single Stage to Orbit (SSTO) with Horizontal Take Off (HTO) using an undercarriage [SSTO(H)];
2. SSTO with HTO using a sled [SSTO-SL];
3. SSTO Air Launched from an aircraft at 8 km altitude, traveling at 180 m/s [SSTO-AL];
4. SSTO with ground-launched Vertical Take Off (VTO) [SSTO(V)];
5. Two Stage To Orbit (TSTO) using an undercarriage [TSTO(H)]; and,
6. TSTO with ground-launched VTO [TSTO(V)].

All stages in all architectures use horizontal landing, and thus have wings and undercarriages. The wings and undercarriages are sized for the vehicle's return, unless they are used in take off, as for example by 1) SSTO(H). All architectures are sized to a common take-off mass (GLOW) of 500 tonnes, but their different architectures result in different payloads delivered to the reference orbit (Table 3).

The primary means of taking advantage of MNT without changing these architectures is to use diamondoid materials with much higher strength-to-density ratios. This will reduce the mass of the vehicles' fuselages, wings and tails, undercarriages, and propulsion systems, roughly in proportion to the relative strength-to-density of diamondoid to the materials used. (Many parts of the propulsion system, however, would require appropriately coated surfaces, and external surfaces will require thermal protection for re-entry.) This research assumes that the systems were all be built of titanium in the original paper. The nature of the "on-board equipment" is not clear described in [5]. It probably includes mass for life support for pilots of the vehicles. Since the exact nature of this mass is unclear, the conservative assumption made here is that this mass could not be reduced at all with the use of MNT.

By flying the same trajectories, the mass savings from applying MNT to the vehicle's structure can be applied to the payload. This occurs directly for the SSTO vehicles and the second stages of the TSTO vehicles. Mass savings on the first stage of the TSTO vehicles are added proportionately to all components of those vehicles second stages, including the payload.

The resulting vehicle and payload masses, and mass ratios, are shown in table 3 below. The "Titanium" based numbers (first entry) are from [5]. The "Diamondoid" based numbers are assuming MNT materials. These results are consistent with [6], which describes an MNT-based SSTO with a payload to GLOW ratio of 1/6 and a ratio of payload mass to dry, empty vehicle mass of ~8.

Table 3. Values for material properties used in this research. The difference from old to new is the use of MNT materials in the place of titanium.

Architecture Titanium/ Diamondoid	"Dry, Empty Vehicle Mass"	Payload Mass	Mass Ratio	"Payload to Dry, Empty Mass Ratio"	Cost per kg to Reference orbit
1) SSTO(H)	92.5 / 16.5	-32 / 44	-6.4% / 8.79%	-35% / 267%	NA / $5.19 / $375
2) SSTO-SL	57.3 / 10.5	9.8 / 56.6	1.96% / 11.33%	17% / 541%	$29k / $3.91 / $185
3) SSTO-AL	54.3 / 9.5	17.0 / 61.8	3.40% / 12.36%	31% / 653%	$16k / $3.54 / $153
4) SSTO(V)	57.4 / 9.4	4.8 / 52.5	0.96% / 10.50%	8% / 540%	$59k / $4.26 / $185
5) TSTO(H)	107.8 / 20.4	17.0 / 49.5	3.40% / 9.90%	16% / 243%	$31k / $4.55 / $412
6) TSTO(V)	70.4 / 14.5	25.0 / 53.5	5.00% / 10.70%	36% / 368%	$14k / $4.17 / $272

The costs for the titanium systems were not in [5], but are based on a very simple minded per mission cost estimate of $1000 per kg of dry, empty launch vehicle mass, averaged over the life of the program. This may be extremely optimistic; using traditional aerospace approaches it would require a tremendous number of fleet missions conducted with low overhead. The cost for the diamondoid, MNT-based systems is based on the high end of the MNT manufacturing cost estimate from , applied to producing the vehicle and the fuel mass, and assuming the vehicle is used once. The third cost entry applies the per mission cost estimate of $1000 per kg of dry, empty launch vehicle mass to the diamondoid system masses.