Status: Study 1991. Payload: 240,000 kg (520,000 lb). Height: 133.00 m (436.00 ft). Diameter: 8.70 m (28.50 ft). Span: 18.00 m (59.00 ft). Apogee: 407 km (252 mi).
The scale of the required launch capability was fundamentally determined by the mass of the payload that will be landed on the Martian surface. The nominal design mass for individual packages to be landed on Mars in the Reference Mission was 50 metric tons for a crew habitat (sized for six people) which must be transferred on a high-energy, fast-transit orbit. This in turn scaled the required mass in LEO to about 240 metric tons.
A number of different technologies could be used to construct a single launch vehicle capable of placing 240 metric tons into a 410-km circular orbit. These launch vehicle concepts used various combinations of past, present, and future U.S. expendable launch vehicle technology and existing launch vehicle technology from Russia and Ukraine.
Option 1used Russian/Ukrainian Energia and Zenit launch vehicle technology combined with STS technology. All of the engines used for this option were existing types that had flown numerous times. The core stage was assumed to be a modification of the existing Energia stage. The modification involved changing the vehicle from one that used a side-mounted payload container to an in-line configuration with strap-on boosters surrounding the core. The upper stage was a new development using STS external tank technology combined with a single SSME. The shroud was entirely new and would be sized for the largest of the Reference Mission payloads. Payload to the assembly orbit was 179 metric tons; this combination of largely existing components did not meet the desired payload launch mass.
Option 2 was a large launch vehicle making extensive use of existing STS and Russian technology. The first stage core and upper stage used the SSME, and the propellant tank structure was based on the STS external tank. Strap-on boosters for this vehicle used the Russian RD-170 engine and a newly designed propellant tank structure. Payload to the assembly orbit was 209 metric tons; this combination also did not meet the desired payload launch mass.
Option 3 used new and old as well as existing technology to create a vehicle that could deliver a payload reasonably close to the desired value. The first stage core propellant tank structure was based on the STS external tank but used newly designed and untested STME engines (the STME program was later cancelled). The strap-on boosters used an updated version of the F-1 engine that powered the first stage of the Saturn V in conjunction with newly designed propellant tanks. The upper stage was comparable to those discussed for the first two options, using STS external tank technology and a single SSME. Payload to the assembly orbit was 226 metric tons.
Option 4 used technology derived from the Saturn V launch vehicle. The first stage core was virtually identical to the first stage of the Saturn V launch vehicle in its basic size and its use of five F-1A engines. Strapped to this stage were four boosters, each with two F-1A engines and roughly one-third of the propellant carried by the core stage. The second stage used six of the J-2 engines that powered the second stage of the Saturn V. However, this upper stage was considerably larger than the Saturn second stage. This last option was the largest of a family of launch vehicles derived using Saturn V launch vehicle technology. Payload to the assembly orbit was 289 metric tons.
The Red Team reviewing DRM v1.0 found Option 3 to be the preferred option. Because a 240-ton-class launch vehicle would be such a development cost issue, consideration was given to the option of launching several hardware elements to LEO using smaller vehicles, assembling (attaching) them in space, and then launching on the outbound trajectory to Mars. This smaller launch vehicle (with a 110- to 120-ton payload capability) would have the advantage of more modest development costs and was in the envelope of capability demonstrated by the unmodified U.S. Saturn V and Russian Energia programs. However, this smaller launch vehicle introduced several potential difficulties to the Reference Mission scenario. The most desirable implementation using this smaller launch vehicle was to simply dock the two elements in Earth orbit and immediately depart for Mars. To avoid boil-off losses in the departure stages (assumed to use liquid hydrogen as the propellant), all elements had to be launched from Earth in quick succession, placing a strain on existing launch facilities and ground operations crews. Assembling the Mars vehicles in orbit and loading them with propellants just prior to departure might alleviate the strain on launch facilities, but the best Earth orbit for Mars missions was different for each launch opportunity, so a permanent construction and/or propellant storage facility in a single Earth orbit introduced additional constraints.
Several launch vehicle designs that could provide this smaller payload capability using existing or near-term technology were examined. The preferred option used the STS external tank for its propellant storage and main structure. Engines for the core stage and the two strap-on boosters were assumed to be the STME engine that was under development at the time of this study.
However the 240-ton payload-class Option 3 launch vehicle was assumed for the Reference Mission. Such a vehicle would require a significant development effort for the launch vehicle, launch facilities, and ground processing facilities; and its cost represented a considerable fraction of the total mission cost. Later versions of the Design Reference Mission would turn to the 75 metric ton payload shuttle-derived Magnum launch vehicle, which was considered less costly to develop.
LEO Payload: 240,000 kg (520,000 lb) to a 407 km orbit at 28.50 degrees.