A3.0 System Design Improvements
Point of Contact: Bret Drake/JSC
In order to accomplish the strategic changes discussed in the previous section, improvements to the system designs were required, specifically in terms of system mass reductions. Modifications to the systems were accomplished by the Exploration Study Team (Johnson Space Center, Marshall Space Flight Center, Lewis Research Center, Ames Research Center, Kennedy Space Center, and Langley Research Center), the JSC TransHab study team, and by the Human/Robotic Exploration Team (led jointly by JPL and JSC). Specific improvements include:
Incorporation of TransHab system designs
Mass scrub of many of the systems
Improvements of the transportation system designs
Point of Contact: Donna Fender/JSC
In an effort to reduce the cost of human habitation in space, a group from the Engineering Directorate at JSC has been studying an economic and innovative habitation concept based on inflatable structure technology. In the spring of 1997, the improvements associated with the TransHab effort were identified as potential habitat options. Many of the subsystem improvements could be incorporated into the Reference Mission for both the interplanetary and surface phases of the mission.
The Exploration Team has been working to quantify improvements identified in the TransHab study, specifically environmental life support system and structural improvements. It is important to note that the Reference Mission architecture and crew size has remained unchanged with the incorporation of the TransHab option. Some of the masses used by the TransHab team, however, have been scaled to match the duration of the Mars Reference Mission.
Advantages of incorporating the TransHab study into the current Mars exploration strategy are manifested primarily in mass reductions. These benefits are provided in Table A3-1. The results are presented for both the Piloted Crew Lander Surface Habitat and the Earth Return Vehicle, and are given in terms of percent changes from the Version 1.0 Reference Mission. Many of the subsystem mass estimates taken from the TransHab studies were of higher fidelity than those previously used by the Mars
Table A3-1 Mass Reduction Benefits from the TransHab Study.
Exploration Team, which accounts for some of the increases in mass values.
Points of Contact: Jerry Sanders and Todd Peters/JSC
The fidelity of the In-Situ Resource Utilization system designs were improved during the Fall 1997 design cycle. An improved system design tool was developed which incorporates options and sizing routines for different products (fuels, oxidizers, water for life support, etc.), production processes, cryogenic fluid cooling, and tank sizing. With the increased fidelity of the model, the ISRU system mass estimates were adjusted downward for the plant itself (from 4802kfrom Reference Mission Version 1.0 to 3941kg) and upward for the hydrogen feedstock (from 4500kg to 5420kg). These estimates reflect a plant that will produce both the ascent propellant and a surface life support system consumables water cache (23 metric tons). The power requirement for the In-Situ Resource Utilization system is driven by both the quantity of products required and the time required to produce the products. Sufficient time for product production is provided such that all required consumables are produced and stored in the surface systems prior to crew departure from Earth. Given these groundrules, the current estimate for the power required is on the order of 45 kWe. Further details describing the mass and power breakdown for the ISRU system are provided in Table A3-2 .
Point of Contact: Bob Cataldo/LeRC
During recent analysis efforts, the surface power system design was revisited in order to obtain mass and cost savings from the original system design. The Reference Mission Version 1.0 surface power system design was based on the reactor technologies developed within the SP-100 program, however with 3-80 kWe closed Brayton cycle (CBC) engines operating at 1100 K. Numerous system trades about this original design were conducted considering power needs, radiation shielding, reactor types, operating temperatures, power conversion technologies, recuperation efficiencies, power distribution voltage, inlet temperature, and number of spare power engines. Updates to the original analysis, including operation at turbine inlet temperatures of 1300 K, enabled a reduction in overall system mass from 14.0 to 10.7 metric tons. Although this assumes a temperature increase of approximately 150 K beyond current Brayton technology, required reactor and fuels technologies remain consistent with those developed within the SP-100 and other DOE/NASA programs. In addition, a first order assessment of the mass impacts of utilizing the same reactor technology as the propulsion system was performed. If feasible and practical, only one development program would then be required for both the propulsion and power systems. A power system based upon a gas-cooled nuclear thermal propulsion engine was estimated to have a mass of
Table A3-2 ISRU System Breakdown for Version 3.0
Table A3-3 Power System Improvements for Version 3.0.
12.1 metric tons. Some of the more salient features of the three designs are shown in Table A3-3. Currently Reference Mission Version 3.0 carries the heavier mass of the gas cooled reactor system.
In addition to the system designs discussed above, other system level trades are being conducted. For instance, additional mass savings could result by using indigenous shielding materials such as soil and/or condensed CO2. The use of indigenous shielding would minimize the system mass differences shown in the table, since the shield mass is the major component of system mass variation.
These concepts are being evaluated for their impact on the power system design itself as well as other systems that might be required to support this concept, such as, mobile equipment or refrigeration systems. In addition, smaller reactor concepts, such as a 50 kWe power system, have been assessed resulting in a total system mass as low as 5.6 metric tons for an SP-100 based system. These smaller reactor concepts could be used for the initial mission phases, with multiple units providing higher power levels for more robust exploration activities, such as food production.
Point of Contact: John Gruener/JSC
A review of the science components for the Reference Mission was conducted during the Fall of 1997. The emphasis of this activity was to critically review the science manifest, seeking mass savings. The focus of the review was not to change the science strategy, but merely to seek methods of reducing the science manifest mass estimates. It is desirable to maintain a balance between mass reduction and science content. Due to the time limitations of the study, it was not possible to conduct detailed system designs for the various scientific instruments, instead, emphasis was placed on understanding the current science content as it pertains to previous systems designs and removing any undefined system content (50 kg), unnecessary undefined margins (250 kg), and undefined discretionary science (300 kg). A detailed science manifest of the first human mission for Reference Mission Version 3.0 is shown in Table A3-4.
Table A3-4 Science Manifest for Version 3.0.
Point of Contact: Robert Yowell/JSC
The EVA consumables estimates for Reference Mission Version 3.0 were improved through the incorporation of a parametric sizing algorithm developed during the TransHab study. In addition, to gain further reductions an assumption was made that consumable mass would only be allocated for two emergency EVAs during transit, allowing for two, eight-hour EVAs, performed by two crew. This resulted in a mass of 48 kg for the transit phases. The transit vehicles also include195 kg each of EVA support equipment (airlock, airlock systems, EMU spares).
System synergism was also incorporated to gain further mass reductions for the surface phase of the mission. EVA consumable requirements were included in the sizing of the In-Situ Resource Utilization system such that additional oxygen was produced by the ISRU system to provide the necessary consumables for routine surface EVA exploration. Utilizing the locally produced oxygen could save approximately five metric tons for the surface phase of the mission alone (10.9 kg per two-person eight-hour EVA). The current EVA consumable estimates are sufficient for one, eight-hour EVA per week, performed by two crew. Additional consumables for a more robust exploration scenario, including food sticks, batteries, drink bags, visors, etc., but not oxygen which the ISRU provides, have not been included in the EVA estimates for Reference Mission 3.0. Estimates for the additional ancillary consumables, for more robust EVAs, will be incorporated in the next version of the Reference Mission. These changes resulted in a total of 446 kg for the surface phase of the mission. Therefore, the total mass of the EVA consumables is currently estimated at 932 kg versus 3000 kg in Version 1.0. Further examination of the assumptions used to reduce these masses is underway.
The EVA dry mass was slightly reduced from 1000 kg to 940 kg based on inputs from the EVA Project Office at Johnson Space Center. This reflects a mass of 156 kg per suit.
Point of Contact: Steve Richards/MSFC
Re-examination of the performance and design characteristics of the transportation elements for the Reference Mission were led by engineers at the Marshall Space Flight Center with support from the Lewis Research Center, Ames Research Center, Langley Research Center, and Kennedy Space Center. Major modifications to the transportation elements, resulting in Reference Mission Version 3.0, are discussed in this section.
A3.3.1 Earth-to-Orbit Transportation
Points of Contact: Bill Eoff and David Smith/MSFC
Human Mars mission launch costs are driven by initial mass in low-Earth-orbit (IMLEO); launch costs per pound of payload; launch vehicle development costs; and on orbit assembly costs. Earth-to-Orbit (ETO) metrics identified in DRM 3.0 required launch vehicle payload capability of 80 metric tons to minimize on orbit assembly costs and meet payload size requirements. Cost metrics of less than $1000 launch cost per pound of payload and total mission costs of $6B for any launch vehicle development costs and all launch recurring costs have been designated as reasonable starting requirements to drive system designs, see Table A3-5.
Table A3-5 Launch Vehicle Requirements
During the design cycle for Reference Mission Version 1.0 numerous configurations were considered and a Shuttle derived vehicle (SDV) with an inline core vehicle was selected. The SDV launch concept barely meets the $6B cost metric for total mission ETO costs because of the high core vehicle costs for Shuttle common hardware. In addition, recent analysis indicated that the SDV configuration exceeded the $1000/lb metric by a factor of two.
Launch vehicle assessments for Reference Mission 3.0 focused on evaluating a core vehicle that is not Shuttle derived to decrease launch costs. Advances in launch vehicle technologies from the Reusable Launch Vehicle (RLV) and Evolved Expendable Launch Vehicle (EELV) programs could make it cost effective to develop a core vehicle that would potentially reduce the $6B ETO cost metric to $2.5B or less per current estimates. This new vehicle concept has been designated as "Magnum" to differentiate from the numerous other past launch vehicle studies. The current Magnum configuration is an inline core vehicle with two attached Shuttle boosters. The payload is aft mounted on the expendable core vehicle; a similar configuration as Titan IV but with over five times the payload capability for one third the launch costs, as shown in Figure A3-1.
Figure A3-1 Payload Capability to 407 km.
The Magnum vehicle configuration includes a core component which is 8.4 meters (27.5 ft) in diameter, the same as the Shuttle External Tank, to allow common use of Shuttle boosters and launch facilities, see Figure A3-2. By using Shuttle launch facilities and the proposed Liquid Fly Back Boosters (LFBB), recurring costs is estimated to be less than $1000 per pound of payload. A composite shroud is used to protect the payload during ascent and a small kick stage is used for circularizing the orbit. The current design of the Magnum launch vehicle provides a delivery capability of 85 metric tons (188 KLB) to 407 km (220 nmi) orbits at 28.5 degrees inclination or 80 metric tons (176 KLB) to 51.6 degree inclination orbits. See Table A3-6 for additional Magnum performance data.
Figure A3-2 Magnum Launch Vehicle.
Technology development and demonstrations for the Magnum launch vehicle concept are driven by the large vehicle size and low life cycle cost requirements. Current evaluations are focused on maximizing the cost-effective application of technologies for engines, valves, composite tanks/structures, and other hardware or facilities under development or projected to be available on other programs such as RLV or EELV. The proposed Magnum technology development program would physically extend these technologies to fit Magnum. Tasks would need to be conducted to demonstrate 8.8 meter (27.5 ft) diameter composite fuel tank manufacturing techniques derived from techniques developed on substantially smaller tanks for RLV. Equivalent
Table A3-6 Payload Capability to 407 km.
tasks would be conducted to demonstrate large composite shrouds using the Advanced Grid Stiffened (AGS) composite shroud manufacturing techniques first developed for EELV by the USAF Phillips Lab. Composite structures, propellant ducts and valve technologies would also need to be demonstrated.
Though the Magnum configuration using LFBBs was selected to drive technology developments, the Magnum configuration is still open for assessment of alternate boosters, engines, etc. which would meet requirements.
Point of Contact: Stan Borowski (LeRC)
A high performance trans-Mars injection (TMI) system is required to propel the cargo and piloted spacecraft payloads from their LEO assembly orbits to the desired trans-Mars trajectories and to stay within the mass (~80 metric tons) and payload dimension (~7.6 m diameter x ~28 m length) limits of the Magnum launch vehicle. For Reference Mission 3.0 the solid core nuclear thermal rocket (NTR) was used for the Trans-Mars Injection stage. Other alternatives, such as a Solar Electric Propulsion concept, are currently under investigation as discussed in Section 5.
Conceptually, the NTR engine is relatively simple (Fig. A3-3). High pressure hydrogen propellant flows from the turbo pumps cooling the nozzle, reactor pressure vessel, neutron reflector, control drums, core support structure and internal radiation shield, and in the process picks up heat to drive the turbines. The hydrogen exhaust is routed through coolant channels in the reactor core’s fuel elements where it absorbs the energy released by fissioning uranium atoms. The propellant is superheated (to 2,700-3,100 K), and then expanded out a supersonic nozzle for thrust. Controlling the NTR engine during its operational phases (startup, full thrust, and shutdown) is accomplished by matching the turbo pump-supplied hydrogen flow to the reactor power level. Control drums, located in the surrounding reflector region, regulate the number of fission-released neutrons that are reflected back into the core. An internal neutron and gamma radiation shield, containing interior coolant passages, is also placed between the reactor core and sensitive engine components to prevent excessive radiation heating and material damage.
Figure A3-3 Schematic of solid core NTR turbopump and power cycle.
The TMI stage used in Reference Mission 3.0 employs three 15 thousand pounds force (klbf) NTR engines, each weighing 2224 kg, for an engine "thrust-to-weight" ratio of ~3.1. The TMI stage utilizes a "tricarbide" fuel material composed of a solid solution of uranium, zirconium and niobium ceramic carbides. This fuel has been developed and extensively tested in Russia. During reactor tests, hydrogen exhaust temperatures of ~3100 K have been reported for run times of over an hour. For exit temperature in the range of 2900-3075 K, specific impulse values of ~940-960 seconds are estimated for the tricarbide NTR engine assuming a chamber pressure of 2000 psia, a nozzle area ratio of 300 to 1, and a 110% bell length nozzle.
A "common" TMI stage design has been defined for both the Mars cargo and piloted missions. The single tank stage is sized for the energetically demanding 2009 fast transit piloted mission opportunity and is therefore capable of injecting heavier surface and orbital payload elements on minimum energy Mars cargo missions. The NTR TMI stage and its aerobraked Mars payloads are illustrated in Fig. A3-4. The TMI stage LH2 tank is cylindrical with …2/2 ellipsoidal domes. It has an inner diameter of 7.4 meters, an ~20 meter length, and a maximum LH2 propellant capacity of ~54 tons assuming a 3% ullage factor. The main TMI stage component is the LH2 tank which is covered by a 2 inch multilayer insulation (MLI) thermal protection system that minimizes propellant boiloff in low Earth orbit to ~0.043 kg/m2/day. Avionics, fuel cell power, storable reaction control system and docking systems are located in the stage forward cylindrical adapter section. Rearward is the stage aft skirt, thrust structure, propellant feed system and NTR engines. The total TMI stage "dry mass" is estimated at ~23.4 metric tons and assumes the use of composite materials for the propellant tank and all primary structures. For the piloted mission, an external disk shield is added to each engine to provide crew radiation protection which increases the stage dry mass by ~3.2 metric tons.
The cargo and piloted Mars spacecraft depart LEO using a "2-perigee burn" Earth departure scenario to reduce gravity losses however single burn departures are also easily accommodated. The total engine burn time for the TMI maneuver is ~35 minutes--about half that demonstrated in the Russian reactor tests. The common TMI stage can inject ~74 and 61 metric tons of payload to Mars on each cargo and piloted mission, respectively. The range of initial mass in Low-Earth Orbit varies from ~135 to 148 metric tons and the overall vehicle length is ~50 meters. Following the TMI maneuver and an appropriate cooldown period, the aerobraked Mars payload and spent TMI stage separate. The storable bipropellant RCS system onboard the TMI stage is then used to perform the final midcourse correction and disposal maneuvers which place the TMI stage onto a trajectory that will not reencounter Earth over the course of a million years.
Points of Contact: Jim Arnold and Paul Wercinski/ARC
The purpose of the Summer/Fall 1997 aeroassist study was to develop and end-to-end conceptual design for human aeroassist vehicles consistent with Reference Mission Version 3.0 payloads and configurations. The emphasis of the study was to develop a reliable mass estimate for the aerobrake as well as to provide a better understanding of the technologies required for the eventual development of an aeroassist capability.
The Aeroassist Summer/Fall study used the Design Reference Mission Version 3.0 Piloted Vehicle mission and trajectory for sizing the entry vehicle for aerocapture and descent from orbit. This trajectory had Mars entry speeds of 7.6 km/s, consistent with a 180-day transit in one particular opportunity. A triconic aerobrake shape was chosen as a baseline to accommodate packaging requirements of the payload elements. It was determined that the triconic shape had sufficient lift-to-drag (L/D) capability to meet aerocapture and descent to surface requirements. An L/D = 0.6 was selected for a trim angle of attack of 47 degrees. The aerocapture at Mars was performed without exceeding the 5g maximum deceleration limit which is necessary to maintain crew health and performance during the aerobraking maneuver.
Figure A3-4 NTR stage and aerobraked Mars payload for Version 3.0
Several Navier-Stokes 3-dimensional flowfield solutions were calculated for this shape using appropriate CO2 chemistry for reacting flows to perform a preliminary thermal protection system (TPS) sizing and trade study for an overshoot trajectory. Turbulent heating estimates were also performed and were identified as a large contributor to uncertainties in predicting heating distribution over the triconic vehicle. Aerodynamic trim was calculated as well and a center-of-gravity location near 49-53% length from the nose was needed for trim. Radiation from the shock layer was also estimated and found to be highly dependent on the reacting gas chemistry models used. Peak heating rates near the nose region were found to between 150-250 W/cm2. Turbulent flows can result in even higher heating rates downstream. For higher entry velocities, at 8.4 km/s, peak heating rates above 350 W/cm2 were modeled, but more analysis is needed due to the higher contributions of radiative heating associated with higher entry speeds. Dust erosion effects were also studied and are expected to not be as large of an effect on TPS mass estimates in comparison to turbulent flow or radiative heating issues. Heatshield structure was only estimated by analogy with structure estimates for a Magnum shroud. Heatshield mass estimates (TPS and
Figure A3-5 Aeroassist Study Results for Version 3.0
structure) yielded mass fractions ranging from 16 – 18% of the total entry vehicle mass. These estimates were used for an entry vehicle carrying 51 metric-tons of cargo. During the aeroassist study, emphasis was not only placed on developing a conceptual approach for human aeroassist, but effort was also devoted to determining key technologies required for aeroassist. The following technology needs the were identified from this study:
Robust 3D Conceptual Fluid Dynamic code capable of radiating, turbulent, and dusty flows
Reliable reacting rate/transport and radiation models
Transition and turbulent models
Validation methods
Guidance Navigation &Control. options on approach, L/D > 0.3 guidance capability, terminal descent and landing
"Human" rated TPS
2D TPS sizing tools
Arc-jets for CO2 flows
Flight validation of TPS materials
High-fidelity integrated design tools supported by local experts across agency.
Points of Contact: Carol Dexter and Larry Kos/MSFC, and Michelle Munk/JSC
Major changes to the descent system for Reference Mission 3.0 include: 1) improved estimates of the descent phase using parachutes, and 2) elimination of the lander mobility requirement.
The descent and landing scheme in Version 1.0 included the use of parachutes with a final landing delta-V of 1000 m/s. The entry to landing phase of the mission was re-examined in Version 3.0 and now includes a higher fidelity method which incorporates mass reductions. Preliminary results were obtained from combining a 3-degree-of-freedom entry simulation and a basic sizing algorithm. In the simulation, the Cargo-1 vehicle, the most massive lander, was deorbited and flown through the atmosphere. Viking-type parachutes were then deployed at about 8 km altitude when the vehicle was traveling roughly 700 m/s. The sizing algorithm was then used parametrically determine the number and size of parachutes and engines required for three different target altitudes. The masses of the parachutes, engines, fuel, and aerobrake were calculated in the sizer and the total vehicle mass was used as the performance metric. The data generated in this analysis are shown in Figure A3-6. A comparison of the new vehicle using the parachute scheme versus the vehicle using the all-propulsive scheme showed a potential savings of ten metric tons. Further analysis of this descent and landing approach includes:
Verifying the results with an integrated simulation
Assessment of supersonic deployment of a cluster of large (on the order of 50-m-diameter) parachutes
Determination of vehicle dynamics
Consideration of aborts, engine-out situations, and hazard avoidance requirements
Reference Mission Version 1.0 included the capability of the descent system to perform limited surface mobility. This capability was provided so that the two surface habitats could be brought together and essentially "docked" to integrate the livable volume for the crew. With the deletion of the initial habitat, the descent system surface mobility mechanisms are not required, thus significantly reducing the complexity of the descent system design.
Figure A3-6 Entry and Landing Parachute Study Results for Version 3.0
Given the improvements in the entry and landing scenario and deletion of the surface mobility requirements, the descent system was refined. The descent system employs four RL10-class engines modified to burn LOX/CH4. These are used to perform the post-aerocapture circularization burn and the final 632 meters per second of descent velocity change after parachute deployment. The descent engines are also used for orbital correction maneuvers during the transit from Earth to Mars, the orbit adjust and trim maneuvers after aerocapture, and the de-orbit burn prior to the atmospheric entry and landing. Architecture definitions for the engine include:
Specific impulse of 379 seconds
Mixture ratio of 3.5
Chamber pressure of approximately 600 psi
A nozzle area ratio of approximately 400
Thrust level of approximately 15,000 lbf.
Additional requirements are that the engines be capable of throttling and gimbaling although specific ranges for these parameters have not been determined.
The descent system for Reference Mission Version 3.0 is capable of placing approximately 40 metric tons of cargo on the surface. The dry mass of this system is approximately 4.9 metric tons requiring 11 metric tons of propellant.
Points of Contact: Carol Dexter and Larry Kos/MSFC
The major modification of the ascent stage for Reference Mission Version 3.0 is the incorporation of a common descent/ascent propulsion system approach. The ascent stage propulsion system shares common engines and propellant feed systems with the descent stage. This eliminates the need for a separate ascent propulsion system reducing the overall mass and subsequent cost. These common engines are the same RL10-class engines modified to burn LOX/ CH4 as the descent stage. These engines perform with an average specific impulse of 379 seconds throughout the ascent maneuver. The ascent propulsion system will require approximately 39 metric tons of propellant to accomplish the approximately 5,625 meters per second of velocity change required for a single-stage ascent to orbit and rendezvous with the previously deployed ERV. The structure and tanks needed for this propellant and the other attached hardware elements have a mass of 4.1 metric tons, including the mass of the engines but not the crew capsule.
Point of Contact: Larry Kos/MSFC
Improvements were also made to the Trans-Earth Injection (TEI) stage for Reference Mission Version 3.0. The TEI stage uses two RL10-class engines modified to burn LOX/ CH4, similar to the descent stage. These engines perform with an average specific impulse of 379 seconds throughout the TEI maneuver. The TEI stage requires approximately 29 metric tons of propellant and has a dry mass of 5.9 metric tons.
Point of Contact: Larry Kos/MSFC
During 1997 the Exploration Transportation Team led my the Marshall Space Flight Center performed a packaging and launch configuration analysis of the Reference Mission Version 3.0 payload elements. The focus of the packaging analysis was to determine the overall launch sequence and payload dimensions to ensure that the mission elements would fit within the overall payload dimensions and launch strategy of the Magnum Launch Vehicle. An overview of the launch packaging analysis is provided in Figure A3-7. As can be seen in the figure, the overall launch sequence of the mission elements begins approximately 97 days prior to the opening of the Trans-Mars Injection window. This timeline is driven primarily by the launch processing of the payload elements and launch vehicle. For this analysis, 30 days were allotted for element processing between launches. A more thorough analysis of the ground processing is currently underway to determine a better estimate for the processing timeline.
Figure A3-7 Launch and Packaging Configurations for Version 3.0
[ Title Page | Table of Contents | Next Section ]