The Mars Reference Mission described in NASA Special Publication 6107, as modified by the updates described in the this addendum, provides a general framework for the human exploration of Mars. Since the original framing of the Reference Mission, other approaches have been brought forward as potential mission and technology options. These approaches, currently being analyzed by the Exploration Team, seem to be promising alternatives for accomplishing the primary objectives set forth in the original mission plan. The major mission alternatives currently under investigation include:
A Solar Electric Propulsion (SEP) option for performing the Earth departure phase of the mission
An approach for capturing the inflated TransHab into Mars orbit
Derivatives of the Nuclear Thermal Rocket concept which produces both propulsive thrust and continuous power
Techniques for minimizing launch mass perhaps to meet a three-Magnum launch scenario, and
All solar power scenarios.
Points of Contact: Kurt Hack and Leon Geffert/LeRC, and Jeff George/JSC
Many different approaches have been developed utilizing both solar electric and nuclear electric propulsion as a method of transporting both cargo and crew to and from Mars. These approaches focused on how an electric vehicle could be utilized to perform all of the major trajectory phases of the mission, including trans-Mars injection, Mars orbit capture, and trans-Earth injection. Although highly efficient from a propellant utilization standpoint, the relatively high power levels required to achieve fast-piloted trips generated two major challenges: 1) The vehicles were very large requiring significant on-orbit assembly and/or deployment, and 2) The technology requirements were significant (lightweight, multi-megawatt-class nuclear or solar powerplants; efficient and durable thrusters scaled to power levels on the order of 500 kWe). These two significant challenges eliminated the electric propulsion vehicle as the primary propulsion concept for human Mars missions.
During the Spring of 1997 an alternate concept of utilizing electric propulsion was proposed by the Lewis Research Center. After examining the payload delivery requirements of the Reference Mission, it was determined that a compromise approach would be to utilize electric propulsion to perform the bulk of the trans-Mars injection, rather than all mission phases. This would minimize the disadvantages of previous approaches while still providing significant mission benefits.
A5.1.1 Electric Propulsion Mission Concept
The solar electric approach currently under investigation by the Exploration Team utilizes the high efficiency of electric propulsion where it provides the most benefit - boosting cargo out of the Earth’s gravity well. The overall mission strategy for the electric propulsion option is fundamentally the same as that of the Reference Mission: two cargo elements are launched in the first mission opportunity, followed by a piloted vehicle in the subsequent opportunity. The only major difference occurs in the replacement of the nuclear thermal TMI stage with a solar electric "tug" and small chemical kick stage.
Injection of cargo and piloted mission elements to Mars begins with the electric propulsion spiral phase. Due to the inherent high specific impulse at low thrust characteristics of electric propulsion, mission elements cannot be directly injected toward Mars via a traditional short impulsive burn. Orbital energy is instead continuously added over a period of approximately nine months, with the vehicle and payload following a spiral trajectory from an initial circular low Earth orbit (LEO) to a final elliptical high Earth orbit (HEO). A small chemical stage is then used to provide the final injection of the mission cargo toward Mars. The now-unloaded solar electric vehicle then returns to LEO to await a repeat sortie of the piloted vehicle element in the succeeding mission opportunity.
Delivery of the crew to Mars requires a slight modification to the front-end of the mission. As with the cargo missions, the electric propulsion vehicle is used to boost the piloted vehicle, sans crew, into a high Earth orbit. The crew is not transported in the vehicle during this phase for two primary reasons. First, during the spiral boost phase of the mission, the vehicle traverses the harsh Van Allen radiation belts many times - far too excessive for piloted missions. Second, the spiral phase takes several months to perform, significantly increasing the exposure of the crew to the debilitating effects of zero-gravity. Rather than employing countermeasures, these effects are minimized by delivering the crew in a high speed taxi to the piloted vehicle after it has been boosted to the final high Earth departure orbit. After a short rendezvous and checkout period, the piloted vehicle, like the previous cargo vehicles, is injected to Mars with a small chemical stage.
A5.1.2 Electric Vehicle Concepts
Vehicle concepts for the electric propulsion option are currently under investigation by the Exploration Team. During the selection and analysis process, emphasis is being placed on developing a concept which can be deployed easily, do not require significant advancements in technology, and is a low cost approach. Conceptual vehicle designs for the crew taxi and solar electric vehicle are shown in Figures A5-2 and A5-3. The concepts shown are still under investigation and will continue to evolve as advancements in the analysis are made. A summary of the mission mass estimates for the solar electric vehicle concept are provided in Figure A5-4.
Figure A5-2 SEP Crew Taxi Concepts.
Point of Contact: Bill Schneider/JSC
The other major mission option currently under investigation is the approach of aerocapturing the inflated TransHab into Mars orbit. During the Fall of 1997 a "Skunk Works" study team composed of experts from the Johnson Space Center, Langley Research Center, Ames Research Center, and Marshall Space Flight Center, conducted a study of the TransHab aerocapture concept. The goals of the study were to design a lightweight aeroshell system capable of capturing the inflated TransHab into Mars orbit and to determine the best system for crew return to Earth. Aeroentry and landing were not considered during this study and were left for follow-on analysis.
Two aeroshell concepts were analyzed during the study: The Ellipsled, which uses the structure from the Magnum launch vehicle shroud (requiring no on-orbit assembly); and the Spherical Dome, which is Shuttle-launched and assembled (see Figure A5-5). In order to estimate the total system mass, the analysis included investigations of the entry flight dynamics, thermal protection system, structural design, and assembly operations.
Analysis conducted showed that both the Ellipsled and the Spherical Dome could accomplish an aerocapture at Mars with positive margins. However, a number of factors has led to a selection of the slender shape concept. Results from the study indicated that the mass fraction (ratio of the mass of the aerobrake to the mass of the aerobrake and payload) ranged from 14.6% to 15.5%. These results closely matched those from the previous aeroshell analyses (see section A3.4), indicating reliability in the mass estimates. Aerobrake mass fractions of this magnitude provide significant mission advantages by reducing the total mass required for the mission.
Analyses of the TransHab Aerobrake concepts are still in progress. Several factors
Figure A5-3 Conceptual Solar Electric Propulsion Vehicle
Figure A5-4 Launch manifest for the Solar Electric Propulsion Vehicle Concept.
remain to be investigated to complete the study, including:
Modifications to the TransHab. The initial effort focused on aerocapturing the habitat originally designed by the JSC team in the Spring of ‘97. Modifications to the TransHab, including structural modifications to operate on the surface of Mars, and the addition of a crew flight deck, were not addressed.
Entry and Landing. The initial study focused only on the aerocapture phase of the mission, and did not address the issues associated with the entry and landing phases, such as static and low-speed dynamic stability, parachute deployment, terminal engine requirements, or landing accuracy. If it is not feasible to land the inflatable TransHab in its current configuration, modifications and additional vehicle elements may have to be introduced into the architecture.
Assessment of the impacts of decreasing the Earth-Mars transit times from 200-days to 180-days to be consistent with previous analysis performed by the Ames Research Center.
TransHab-to-Aeroshell Attachment. Detailed structural design of the attachments between the TransHab and the aeroshell will serve to further refine the mass estimates of the system, and must address packaging, deployment, and accessibility issues.
Figure A5-5 Potential TransHab Aerobrake Configurations.
Point of Contact: Stan Borowski/LeRC
Although most of the current work is focused on the Solar Electric concept, the NTR approach is being maintained for comparison. The solid core NTR propulsion system represents a "rich source of energy" in that it contains substantially more uranium-235 fuel in its reactor core than it consumes during its primary propulsion maneuvers. By configuring the NTR engine as a "bimodal" system, abundant electrical power can also be generated for a variety of spacecraft needs. During power generation, the reactor core operates in essentially an "idle mode" with a thermal power output of ~100 kilowatts. The reactor thermal energy is subsequently removed and routed to a turbo-alternator-compressor Brayton power conversion unit using a helium-xenon working fluid, as shown in Figure A5-6. A space radiator system rejects waste heat and also reduces decay heat propellant loss following propulsive burns.
Figure A5-6 Schematic "Bimodal" NTR System
5.3.1 Bimodal NTR Mission Concept
An option to Reference Mission Version 3.0 that utilizes bimodal NTR transfer vehicles in place of the expendable NTR stages is being evaluated. A common "core" stage, used on cargo and piloted vehicles alike, is outfitted with three 15 klbf bimodal NTR engines capable of providing up to 50 kilowatts of electrical power (kWe) using any two engines The bimodal core stage is not jettisoned after the TMI maneuver but remains with the cargo and piloted payload elements providing midcourse correction (MCC) propulsion and all necessary power during transit. Near Mars, the bimodal stage separates from the aerobraked payloads and performs its final disposal maneuvers. A key difference between Reference Mission 3.0 and the bimodal option is the absence of the aerobraked LOX/methane (CH4) TEI stage which is replaced by an "all propulsive" bimodal NTR-powered Earth Return Vehicle (ERV) illustrated in Figure A5-7.
The bimodal stage LH2 tank is slightly shorter than the expendable TMI stage tank at 19 meters and has a maximum LH2 propellant capacity of ~51 tons with a 3% ullage factor. A turbo-Brayton refrigeration system is located in the forward cylindrical adaptor section to eliminate LH2 boiloff during the lengthy (~4.3 year) ERV mission. A 12 kWe Brayton refrigeration system is included to remove the ~100 watts of heat flux penetrating the 2 inch MLI system in low-Earth-orbit where the highest heat flux occurs. Enclosed within the conical aft radiator section of the bimodal core stage is a closed Brayton cycle (CBC) power conversion system employing three 25 kWe Brayton rotating units (one for each bimodal reactor) which operate at ~2/3 of rated capacity, thus providing an "engine out" capability. The turbine inlet temperature of the He-Xe working gas is ~1300 K and the total system specific mass is estimated to be ~30 kg/kWe.
A mass comparison of the bimodal NTR transfer
Figure A5-7 "Bimodal" NTR Transfer Vehicle Option
Table A5-1 Comparison of "Bimodal" NTR to Reference Mission Version 3.0
vehicles and the Reference Mission Version 3.0 vehicles is shown in Table A5-1.
The mass values assume a "2-perigee burn" Earth departure scenario. Overall, the bimodal approach has a lower "three-mission" initial mass than Reference Mission 3.0. In addition, the bimodal approach can reduce the operational complexity of the mission (eliminates solar array deployment/retraction) as well as eliminating the need for an aerobrake and injection stage for the Earth Return Vehicle.
5.3.2 All Propulsive" Bimodal NTR Option Using TransHab
Another option to the Reference Mission 3.0 under consideration is the use of a bimodal NTR stage to propulsively capture all payload elements into Mars orbit. This "all propulsive" NTR option provides the most efficient use of the bimodal engines which can supply abundant power to the spacecraft and payloads in Mars orbit for long periods. Propulsive capture into the reference "250 km by 1 sol" elliptical Mars parking orbit also makes possible the use of a standardized, reduced mass "aerodescent" shell because of the lower payload entry velocity (~4.5 km/s) encountered. From this orbit, the triconic aerobrake mass varies by only ~400 kg for a 20 ton increase in payload mass (see Section 3.3.3).
The attractiveness of the "all propulsive" bimodal NTR option is further increased by the utilization of the lightweight, inflatable "TransHab" module discussed in Section 3.1. The substitution of TransHab for the heavier, hard-shell habitat module introduces the potential for propulsive recovery of the Earth Return Vehicle in Earth orbit and its reuse on subsequent missions. TransHab use also allows the crew to travel to and from Mars on the same bimodal transfer. In Mars orbit, the crew transfer vehicle rendezvous with the "unpiloted" habitat lander which is now delivered as a cargo element by the bimodal stage. The absence of crew from the bimodal habitat lander eliminates the need for outbound consumables and engine crew radiation shields and allows it to carry off-loaded surface habitation and science equipment previously carried on the cargo lander.
A three-dimensional image of the bimodal transfer vehicle used on the piloted mission is shown in Figure A5-8. The TransHab is ~9.7 meters long and inflates to a diameter of ~9.5 meters. Its total mass is ~24.3 metric tons which includes the crew and their consumables. The total length and initial mass of the piloted transfer vehicle is ~54 meters and ~141 metric tons, respectively. A smaller, "in-line" propellant tank is used on the bimodal transfer vehicles that deliver the ~46 metric ton habitat and ~54 ton cargo landers into Mars orbit. The habitat and cargo transfer vehicles are ~56 meters long and have a LEO mass of ~129 metric tons and 144 metric tons, respectively.
5.3.3 "LOX-Augmented" NTR Option
An enhanced NTR option, known as the "LOX-augmented" NTR (LANTR), is presently under study by NASA which combines conventional LH2-cooled NTR and supersonic combustion ramjet (scramjet) technologies. The LANTR concept utilizes the large divergent section of the NTR nozzle as an "afterburner" into which LOX is injected and supersonically combusted with reactor preheated hydrogen emerging from LANTR’s choked sonic throat--essentially "scramjet propulsion in reverse." By varying the oxygen-to-hydrogen mixture ratio (MR), the LANTR engine could potentially operate over a wide range of thrust and specific impulse values while the reactor core power level remains relatively constant. For those missions where volume (not mass) constraints limit bimodal stage performance, LANTR propulsion can help to increase "bulk" propellant density and total thrust output, while decreasing the engine burn times. LOX augmentation would be particularly beneficial during the TMI burn to reduce gravity losses. Following this maneuver, the spent "in-line"
LH2 tank and a small LOX tank attached to it could be jettisoned as a single unit. On all subsequent burns, the LANTR engines would operate on only LH2 (MR = 0). Cold flow experimental injector tests and reactive computational fluid dynamics analyses are currently underway at NASA Lewis Research Center in preparation for future hot flow tests aimed at demonstrating concept feasibility.
Figure A5-8 All Propulsive Bimodal NTR Carrying TransHab
Point of Contact: Andrew Petro/JSC
During the Spring of 1998 a special design study was conducted to define the elements, mission content, and technology required to accomplish a human Mars mission which could be accommodated for launch within the mass and volume capacity of three heavy-lift launch vehicles. The reference launch vehicle used in the study was the Magnum launch vehicle and so this mission concept is referred to as the "Three-Magnum Mars Mission". The design team was directed to employ a solar electric propulsion (SEP) stage for delivering the Mars mission elements to a high apogee Earth departure orbit and to not employ nuclear propulsion for any maneuvers.
This study was unusual in the approach of designing to a fixed constraint for Earth launch mass. The most significant result of the study was the identification of the technology challenges which must be met to achieve the launch mass goal.
The capacity of the Magnum launch vehicle defined for this study was 89.5 metric tons for launch packages which employ the launch shroud as an aeroshell, and 85.5 metric tons for payloads which do not include the shroud as payload. The payload capability quoted is for launch from the Kennedy Space Center to a circular orbit of 400 kilometers at an inclination of 28.5 degrees. The dimensions of the Magnum shroud were defined as an outer diameter of 8.4 meters and a length of 28 meters.
The mission defined in this study included a crew of four people, a scientific payload of 1770 kg and two unpressurized rovers with a mass of 650 kg each. The missions were conjunction-class with outbound and inbound transit durations of 180 to 200 days and Mars surface stay times of 520 to 580 days. The elements were designed to accomplish missions in six out of the eight opportunities in the synodic cycle. The other two opportunities would require an additional propulsive stage of approximately 16 metric tons.
Several different mission scenarios were considered and two were documented for the study: a Combination Lander Scenario in which all elements are sent to Mars in a single opportunity, and a Split Mission Scenario in which some elements are deployed at Mars in the first opportunity and the crew travels to Mars in the next opportunity. The Split Mission Scenario is similar to the Design Reference Mission 3.0 whereby propellant for Mars ascent is produced at Mars.
5.3.4.2 Strategies and Technology Challenges
Several strategies were used to constrain the total mission mass with respect to the Design Reference Mission and to achieve the launch mass target.
Crew reduced from 6 to 4 persons
Initial departure orbit apogee raised from 39,000 km to 120,000 km
Hydrogen fuel is used for all maneuvers.
In addition, several technology development challenges were identified as necessary to achieve the launch mass target.
Structures, tanks, and aeroshells with a reduction in mass of up to 50% over current technology
High performance power generation system for space and surface operations (100 kg/kWe)
Long-term hydrogen storage with near zero boil-off for up to four years
Lightweight chemical propulsion engines with a specific impulse of 480 sec.
Deployable solar electric propulsion system with a megawatt-capacity solar array
5.3.4.3 Combination Lander Scenario
This scenario is illustrated in Figure A5-9 and the launch packages with element masses are shown in Figure A5-10. Figure A5-11 is a three-dimensional drawing of the Combination lander concept as it would be deployed on the surface. This lander includes the crew module for descent and ascent along with the surface habitat.
5.3.4.4 Split Mission Scenario
The Split Mission Scenario is similar to the Design Reference Mission scenario but it includes all of the strategies and technology challenges mentioned above. The major differences in this scenario are 1) the pre-deployment of the return vehicle in Mars orbit, 2) pre-deployment of the ascent vehicle on the surface of Mars, 3) the production of propellant on Mars, and 4) the use of methane rather than hydrogen for Mars ascent. The scenario is illustrated in Figure A5-12 and the launch packages and element masses are shown in Figure A5-13.
5.3.4.5 Summary
By incorporating the aggressive technology goals, two mission scenarios were defined which could meet the three-Magnum launch mass and volume constraint. It should be noted that each scenario also requires a Space Shuttle launch at the beginning of the mission to deliver the crew and their high-Earth orbit taxi and also a Shuttle mission at the end to recover the crew in low Earth orbit. This three-launch strategy is reliant on the key technologies described previously. An effort is currently underway to better understand the difficulty of the technology challenges as they compare to current state-of-the-art, the risks associated with these technologies, development costs, and the architectural impacts of potential technology fall-backs if it is believed that the technology development cannot be completed as needed.
Figure A5-9 Three-Magnum Combination Lander Scenario
Figure A5-10 Three-Magnum Combination Lander Launch Packages
Figure A5-11 Combination Lander Concept on Mars Surface
Figure A5-12 Three-Magnum Split Mission Scenario
Figure A5-13 Three-Magnum Split Mission Launch Packages
Another alternative strategy under consideration is an approach where total reliance would be place on propulsion and power concepts based solely on chemical and solar technologies. Of particular importance is the power generation strategy which has relied on the same technology base (SP-100) from the original reference through Reference Mission 3.0. This power strategy has been one of providing a robust power generation and storage capability to enable significant mass reductions. This technique of trading mass for power has been manifested in the Reference Mission in the form of advanced technologies such as in-situ resource utilization, bioregenerative closed-loop life support systems, and long-range pressurized rovers. These high power demands necessitated the use of advanced power concepts such as surface nuclear reactors and dynamic isotope power sources.
A major challenge of an all-solar human Mars mission is the lack of solar irradiation at Mars. As can be seen in Figure 5-14 the solar flux at the surface can be as low as 6.5% of that in low-Earth orbit. The reduction of solar flux is due to the distance of Mars from the sun, the presence of the atmosphere, and potential dust storms.
Figure 5-14 Solar Irradiation At Mars.
Analysis of an all-solar approach will include:
Developing a mission approach where the surface element power needs can be reduced to the lowest level possible
Understanding the sensitivities of advanced solar cell technologies
Analysis of solar power generation system setup and maintenance, such as cleaning due to dust accumulation
Analysis of the impacts of elimination of advanced technologies (in-situ resource utilization, long-range rovers, food production, etc.) on the overall mission approach, including risk.
Analysis of this all-solar mission approach is currently under way. Results of this study will be included in the next update of the Reference Mission.
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