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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1996

Refinement of planetary protection policy for Mars missions

Under existing COSPAR policy adopted in 1984, missions to Mars (landers, probes, and some orbiters) are designated as Category IV missions. As such, the procedures for implementing planetary protection requirements could include trajectory biasing, cleanrooms, bioload reduction, sterilization of hardware, and bioshields. In 1992, a U.S. National Research Council study recommended that controls on forward contamination of Mars be tied to specific mission objectives. The report recommended that Mars landers with life detection instruments be subject to at least Viking-level sterilization procedures for bioload reduction, while spacecraft (including orbiters) without life detection instruments be subject to at least Viking-level pre-sterilization procedures for bioload reduction but need not be sterilized. In light of this, it is proposed that the current policy's Category IV and its planetary protection requirements be divided into two sub-categories as follows: Category IVa, for missions comprising landers and probes without life detection experiments, which will meet a specified bioburden limit for exposed surfaces, and Category IVb, for landers and probes with life detection experiments, which will require sterilization of landed systems. In addition, Category III orbiter mission specifications are expanded to be consistent with these recommendations.

NASA Discipline Exobiology↗
NASA NTRS2023

Parametric Modeling of NTP Engine Performance for a Crewed Mars Mission

Nuclear thermal propulsion (NTP) has been studied extensively for use on a crewed mission to Mars, and NASA and DARPA are partnering in the development and demonstration of an NTP engine as part of the DRACO project. The specific impulse and thrust achieved by an engine in the context of a mission determines the propellant amount required to close the mission and the size of the vehicle required. An NTP engine has a slower start-up/shut-down transient than traditional chemical engines and requires post-firing cooling, and thus the average Isp for a burn is a function of the power ramp-up rate and the duration of the burn. An initial demonstration engine may target a lower Isp than the 900 sec that has been extensively studied by NASA. This paper shows the effect on the mission closure and vehicle sizing of the transient ramp rate performance and of lower than 900 sec specific impulse. Ammonia is considered as an alternate NTP propellant, with mission performance estimated.

NTP↗
NASA NTRS1994

Revision to Planetary Protection Policy for Mars Missions

Under existing COSPAR policy adopted in 1984, missions to Mars (landers, probes, and some orbiters) are designated as Category IV missions. As such, the procedures for implementing planetary protection requirements could include trajectory biasing, cleanrooms, bioload reduction, sterilization of hardware, and bioshields, i. e. requirements could be similar to Viking. However, in 1992, a U. S. National Academy of Sciences study recommended that controls on forward contamination of Mars be tied to specific mission objectives. The report recommended that Mars landers with life detection instruments be subject to at least Viking-level sterilization procedures for bioload reduction, while spacecraft (including orbiters) without life detection instruments be subject to at least Viking-level pre sterilization procedures for bioload reduction but need not be sterilized. In light of this, it is proposed that the current policy's Category IV missions and their planetary protection requirements be divided into two subcategories as follows: Category IV A, for missions comprising landers and probes without life detection experiments and some orbiters, which will meet a specified bioburden limit for exposed surfaces; Category IV B, for landers and probes with life detection experiments, which will require complete system sterilization. For Category IV A missions, bioburden specifications will be proposed and implementing procedures discussed. A resolution will be proposed to modify the existing COSPAR policy to reflect these changes. Similar specifications, procedures, and resolution for Category IV B missions will be the subject of a later study.

DeVincenzi, D. L.↗
NASA NTRS1991

Mars Observer - The next Mars mission

The next mission to Mars, called Mars Observer, will be launched in September 1992. After the capture of the spacecraft by the planet and the adjustment into a low, sun-synchronous, polar-mapping orbit in late 1993, observations will continue for a Mars year (687 days). The scientific mission centers around global geoscience and climatology observations of the Mars atmosphere, surface, and interior. The seven experiments carried by the spacecraft involve gamma-ray spectroscopy, magnetometry, surface and atmospheric imaging, atmospheric sounding, laser altimetry, gravity mapping, and thermal emission spectroscopy. All experiments contain microprocessors, which will be controlled remotely from the investigator's home institution. The long planned period of continuous 24 h/day observation promises a rich harvest of global and seasonal information. Mars Observer stands between the initial exploration of Mars and the more intensive explorations, possibly involving human beings, that are only now being planned.

Albee, Arden L.↗
NASA NTRS1992

Physiologically constrained aerocapture for manned Mars missions

Aerobraking has been proposed as a critical technology for manned missions to Mars. The variety of mission architectures currently under consideration presents aerobrake designers with an enormous range of potential entry scenarios. Two of the most important considerations in the design of an aerobrake are the required control authority (lift-to-drag ratio) and the aerothermal environment which the vehicle will encounter. Therefore, this study examined the entry corridor width and stagnation-point heating rate and load for the entire range of probable entry velocities, lift-to-drag ratios, and ballistic coefficients for capture at both Earth and Mars. To accomplish this, a peak deceleration limit for the aerocapture maneuvers had to be established. Previous studies had used a variety of load limits without adequate proof of their validity. Existing physiological and space flight data were examined, and it was concluded that a deceleration limit of 5 G was appropriate. When this load limit was applied, numerical studies showed that an aerobrake with an L/D of 0.3 could provide an entry corridor width of at least 1 degree for all Mars aerocaptures considered with entry velocities up to 9 km/s. If 10 km/s entries are required, an L/D of 0.4 to 0.5 would be necessary to maintain a corridor width of at least 1 degree. For Earth return aerocapture, a vehicle with an L/D of 0.4 to 0.5 was found to provide a corridor width of 0.7 degree or more for all entry velocities up to 14.5 km/s. Aerodynamic convective heating calculations were performed assuming a fully catalytic, 'cold' wall; radiative heating was calculated assuming that the shock layer was in thermochemical equilibrium. Heating rates were low enough for selected entries at Mars that a radiatively cooled thermal protection system might be feasible, although an ablative material would be required for most scenarios. Earth return heating rates were generally more severe than those encountered by the Apollo vehicles, and would require ablative heat shields in all cases.

Lyne, James Evans↗
NASA NTRSDate not supplied

Nuclear Electric Propulsion (NEP) Developments for Human Mars Missions

Nuclear Electric Propulsion (NEP) has the potential to enable human missions to Mars and other ambitious missions throughout the solar system. This paper will discuss ongoing NEP research and development efforts and challenges to developing NEP systems for these missions. The baseline NEP vehicle design is comprised of multiple subsystems operating in concert. The major subsystems include a nuclear reactor, a dynamic power conversion system to generate electrical power, heat rejection/radiators, a power management and distribution system to deliver power to the electric propulsion (EP) system and to the crewed transit habitat, and the EP system. An additional chemical propulsion system is also included to execute higher-thrust maneuvers in the Earth and Martian gravity wells. Specific issues that will be addressed include mission trajectory design and optimization, concurrent subsystem technology advancement plans and maturation efforts, and systems-level modeling of the entire vehicle to permit the fast exploration of trade-offs that can occur with changes in operating conditions or through the use of different technologies.

Kurt A Polzin↗
NASA NTRS2013

A One-year, Short-Stay Crewed Mars Mission Using Bimodal Nuclear Thermal Electric Propulsion (BNTEP) - A Preliminary Assessment

A crewed mission to Mars poses a significant challenge in dealing with the physiological issues that arise with the crew being exposed to a near zero-gravity environment as well as significant solar and galactic radiation for such a long duration. While long surface stay missions exceeding 500 days are the ultimate goal for human Mars exploration, short round trip, short surface stay missions could be an important intermediate step that would allow NASA to demonstrate technology as well as study the physiological effects on the crew. However, for a 1-year round trip mission, the outbound and inbound hyperbolic velocity at Earth and Mars can be very large resulting in a significant propellant requirement for a high thrust system like Nuclear Thermal Propulsion (NTP). Similarly, a low thrust Nuclear Electric Propulsion (NEP) system requires high electrical power levels (10 megawatts electric (MWe) or more), plus advanced power conversion technology to achieve the lower specific mass values needed for such a mission. A Bimodal Nuclear Thermal Electric Propulsion (BNTEP) system is examined here that uses three high thrust Bimodal Nuclear Thermal Rocket (BNTR) engines allowing short departure and capture maneuvers. The engines also generate electrical power that drives a low thrust Electric Propulsion (EP) system used for efficient interplanetary transit. This combined system can help reduce the total launch mass, system and operational requirements that would otherwise be required for equivalent NEP or Solar Electric Propulsion (SEP) mission. The BNTEP system is a hybrid propulsion concept where the BNTR reactors operate in two separate modes. During high-thrust mode operation, each BNTR provides 10's of kilo-Newtons of thrust at reasonably high specific impulse (Isp) of 900 seconds for impulsive transplanetary injection and orbital insertion maneuvers. When in power generation/EP mode, the BNTR reactors are coupled to a Brayton power conversion system allowing each reactor to generate 100's of kWe of electrical power to a very high Isp (3000 s) EP thruster system for sustained vehicle acceleration and deceleration in heliocentric space.

Burke, Laura M.↗
NASA NTRS2005

Technology Development for NASA Mars Missions

A viewgraph presentation on technology development for NASA Mars Missions is shown. The topics include: 1) Mars mission roadmaps; 2) Focus and Base Technology programs; 3) Technology Infusion; and 4) Feed Forward to Future Missions.

Hayati, Samad↗
NASA NTRS2013

A One-year, Short-Stay Crewed Mars Mission Using Bimodal Nuclear Thermal Electric Propulsion (BNTEP) - A Preliminary Assessment

A crewed mission to Mars poses a signi cant challenge in dealing with the physiolog- ical issues that arise with the crew being exposed to a near zero-gravity environment as well as signi cant solar and galactic radiation for such a long duration. While long sur- face stay missions exceeding 500 days are the ultimate goal for human Mars exploration, short round trip, short surface stay missions could be an important intermediate step that would allow NASA to demonstrate technology as well as study the physiological e ects on the crew. However, for a 1-year round trip mission, the outbound and inbound hy- perbolic velocity at Earth and Mars can be very large resulting in a signi cant propellant requirement for a high thrust system like Nuclear Thermal Propulsion (NTP). Similarly, a low thrust Nuclear Electric Propulsion (NEP) system requires high electrical power lev- els (10 megawatts electric (MWe) or more), plus advanced power conversion technology to achieve the lower speci c mass values needed for such a mission. A Bimodal Nuclear Thermal Electric Propulsion (BNTEP) system is examined here that uses three high thrust Bimodal Nuclear Thermal Rocket (BNTR) engines allowing short departure and capture maneuvers. The engines also generate electrical power that drives a low thrust Electric Propulsion (EP) system used for ecient interplanetary transit. This combined system can help reduce the total launch mass, system and operational requirements that would otherwise be required for equivalent NEP or Solar Electric Propulsion (SEP) mission. The BNTEP system is a hybrid propulsion concept where the BNTR reactors operate in two separate modes. During high-thrust mode operation, each BNTR provides 10's of kilo- Newtons of thrust at reasonably high speci c impulse (Isp) of 900 seconds for impulsive trans-planetary injection and orbital insertion maneuvers. When in power generation / EP mode, the BNTR reactors are coupled to a Brayton power conversion system allowing each reactor to generate 100's of kWe of electrical power to a very high Isp (3000 s) EP thruster system for sustained vehicle acceleration and deceleration in heliocentric space.

Burke, Laura A.↗
NASA NTRS1979

Atmospheric entry systems for advanced Mars missions

A study of estimates of the mass and performance characteristics of entry system designs for advanced Mars missions is presented. The missions were: (1) a Mars Surface Sample Return Mission (MSSR), (2) a Mars Hard Lander, and (3) a Mars Aircraft Mission, and Viking technology or its extensions were used for these entry system conceptual designs. For the MSSR mission, various mission combinations were evaluated including different Mars Ascent Vehicle masses and the use of 2 through 4 stages of the IUS launch vehicle. The constraints for this Mars mission were the Shuttle payload bay geometry limitation and its massive requirements. Requirements for the simple Hard Lander Mission including entry and deceleration, using a parachute and a solid rocket are discussed. A design requirement for the Mars Aircraft Mission to minimize the total entry capsule/aircraft mass by combining entry functions and hardware into the aircraft system was described, and various approaches for supporting the folded aircraft within the aeroshell were examined.

Butts, A. J.↗
NASA NTRS1986

Manned Mars mission communication and data management systems

A manned Mars mission will involve a small crew and many complex tasks. The productivity of the crew and the entire mission will depend significantly on effective automation of these tasks and the ease with which the crew can interface with them. The technology to support a manned Mars mission is available today; however, evolving software and electronic technology are enabling many interesting possibilities for increasing productivity and safety while reducing life cycle cost. Some of these advanced technologies are identified.

White, Ronald E.↗
NASA NTRS1991

Nuclear thermal rocket propulsion application to Mars missions

Options for vehicle configurations are reviewed in which nuclear thermal rocket (NTR) propulsion is used for a reference mission to Mars. The scenario assumes an opposition-class Mars transfer trajectory, a 435-day mission, and the use of a single nuclear engine with 75,000 lbs of thrust. Engine parameters are examined by calculating mission variables for a range of specific impulses and thrust/weight ratios. The reference mission is found to have optimal values of 925 s for the specific impulse and thrust/weight ratios of 4.0 and 0.06 for the engine and total stage ratios respectively. When the engine thrust/weight ratio is at least 4/1 the most critical engine parameter is engine specific impulse for reducing overall stage weight. In the context of this trans-Mars three-burn maneuver the NTR engine with an expander engine cycle is considered a more effective alternative than chemical/aerobrake and other propulsion options.

Emrich, W. J., Jr.↗
NASA NTRS1986

Manned Mars Missions. Working group papers, volume 1, section 1-4

The papers presented by the working group on Manned Mars Missions are given. The purpose is to update earlier Mars missions study data, to examine the impact of new and emerging technologies on Mars mission capabilities, and to identify technological issues that would be useful in projecting scientific and engineering research in the coming decades. The papers are grouped into nine sections, which are: (1) rationale; (2) transportation trades and issues; (3) mission and configuration concepts; (4) surface infrastructure; (5) science investigations and issues; (6) life science/medical issues; (7) subsystems and technology development requirements; (8) political and economic issues; and (9) impact on other programs.

Michael B Duke↗
NASA NTRS2003

Life support approaches for Mars missions

Life support approaches for Mars missions are evaluated using an equivalent system mass (ESM) approach, in which all significant costs are converted into mass units. The best approach, as defined by the lowest mission ESM, depends on several mission parameters, notably duration, environment and consequent infrastructure costs, and crew size, as well as the characteristics of the technologies which are available. Generally, for the missions under consideration, physicochemical regeneration is most cost effective. However, bioregeneration is likely to be of use for producing salad crops for any mission, for producing staple crops for medium duration missions, and for most food, air and water regeneration for long missions (durations of a decade). Potential applications of in situ resource utilization need to be considered further. c2002 Published by Elsevier Science Ltd on behalf of COSPAR.

NASA Discipline Life Support Systems↗
NASA NTRS1993

Magnesium and Carbon Dioxide - A Rocket Propellant for Mars Missions

A rocket engine for Mars missions is proposed that could utilize CO2 accumulated from the Martian atmosphere as an oxidizer. For use as possible fuel, various metals, their hydrides, and mixtures with hydrogen compounds are considered. Thermodynamic calculations show that beryllium fuels ensure the most impulse but poor inflammability of Be and high toxicity of its compounds put obstacles to their applications. Analysis of the engine performance for other metals together with the parameters of ignition and combustion show that magnesium seems to be the most promising fuel. Ballistic estimates imply that a hopper with the chemical rocket engine on Mg + CO2 propellant could be readily developed. This vehicle would be able to carry out 2-3 ballistic flights on Mars before the final ascent to orbit.

Shafirovich, E. IA.↗
NASA NTRS2000

Artificial Gravity for Mars Missions: The Different Design and Development Options

One of the major impediments to human Mars missions is the development of appropriate countermeasures for long term physiological response to the micro-gravity environment. A plethora of countermeasure approaches have been advanced from strictly pharmacological measures to large diameter rotating spacecraft that would simulate a 1-g environment (the latter being the most conservative from a human health perspective). The different approaches have significantly different implications not only on the overall system design of a Mars Mission Vehicle (MMV) but on the necessary earth-orbiting platform that would be required to qualify the particular countermeasure system. it is found that these different design options can be conveniently categorized in terms of the order of magnitude of the rotation diameter required (100's, 10's, 1's, 0 meters). From this, the different mass penalties associated with each category can be generally compared. The overall objective of the countermeasure system should be to maximize crew safety and comfort, minimize exercise protocol time (i.e., the time per day that each crew member would have to participate in the exercise/countermeasure), maximize countermeasure effectiveness, and minimize the associated system mass penalty of the Mars Mission Vehicle (in terms of fraction of IMLEO - Injected Mass in Low Earth Orbit).

Murbach, Marcus↗