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Enabling Multiple Abort Strategies Using the NTP Approach for Human Mars Missions
Abstract
Future human exploration missions to Mars are being studied by NASA and industry. Several approaches to the Mars mission are being examined that use various types of propulsion for the different phases of the mission. The choice and implementation of propulsion system options can significantly impact mission performance in terms of trip time, spacecraft mass, and mission abort capability for the crew. Understanding the trajectory requirements relative to the round-trip Earth to Mars mission opportunities in the 2030’s, and beyond, that set the design of the Mars crew vehicle is important in order to determine the impact of key propulsion choices. Additionally some propulsion and propellant choices for the crew vehicle can enable mission abort trajectories while others will not.
Nuclear Thermal propulsion (NTP) with 900 seconds and higher specific impulse (Isp) can permit the capability to abort back to Earth at several opportunities during the outbound or return trajectory that regular low Isp chemical propulsion cannot. Abort scenarios after the Mars crew vehicle has been injected along the path to Mars have been studied as well as timing of fly-by aborts to quickly return crew to Earth.
These trajectory studies are based off missions NASA defined during the Evolvable Mars Campaign (EMC) with crew going to Mars in 2033, 2037, 2043 and 2048. Detailed trajectory analysis was performed with the NASA Copernicus program for the several crew missions that were in the EMC. The goal was to determine how the heliocentric trajectory elements change and determine the “abort trajectory” impulse requirements. The NTP Mars crew vehicles studied had propellant loads sized for either Trans-Mars Injection (TMI), Mars Orbit Capture (MOC), Trans-Earth Injection (TEI), and Earth Orbit Capture (EOC) or just TMI and MOC with TEI and EOC pre-positioned as was done in the EMC chemical crew vehicle architecture.
The impulse requirements were then compared to the remaining impulse capability after TMI for the architecture’s remaining propellant load for the baseline 2033, 2037, 2043, and 2048 crew vehicle concepts. Abort scenarios that were studied included fast returns some number of days after TMI as well as fly-by aborts using available propellants (e.g., main propulsion system (MPS) and reaction control system (RCS)). The NTP system in this study shows great promise for adding mission flexibility via abort trajectories to Mars and at Mars.
... Since 2019 other fuel materials have been studied such as CERCER (Ceramic-Ceramic) (e.g., Uranium Nitride-Zirconium Carbide). The design process for HALEU NTP studies has included detailed thermodynamic cycle analysis and has been documented in several published papers [2,3,4,5,6]. Fig. 1 [19] shows a simplified rocket thermal cycle; specifically, an expander cycle configuration suitable for the 25,000 lbf thrust NTP. ...
... Its fundamental performance capabilities are double that of current, state-of-the-art technologies, making it an optimal solution for the high mass, long distance architectures necessary for human spaceflight beyond cislunar space. NTP is the only currently viable option that delivers six astronauts to Mars within three months, reduces the total required number of launches, enables mission profiles with either short-term or extended stays on the Martian surface, and offers abort scenarios at any point during transit [1][2][3][4] . ...
This paper establishes the feasibility of Low-Enriched Uranium fueled Nuclear Thermal Propulsion (LEU-NTP) reactors in the low-thrust region below 16klbf (71kN). A reference, 7.5klbf, High-Enriched Uranium (HEU) design is converted to 19.75% enriched LEU and shown to be capable of reaching criticality and meeting other performance requirements with a minimal mass increase. At this smaller scale, historical LEU-NTP conversion techniques that focus on maximizing neutron moderation or minimizing leakage within the active core are insufficient. Thus, the 7.5klbf preliminary design requires several unique modifications, including a reduction in the number of control drums and adjustments to selected materials. To verify the design’s feasibility, several key neutronic and thermal-hydraulic performance parameters including burnup, Xenon worth, and submersion criticality are characterized with Serpent 2.0 and the Space Propulsion Optimization Code. The methods applied in this work reveal an opportunity for further LEU-NTP optimization that may directly translate to increased scalability and efficiency for future designs.
... Its fundamental performance capabilities are double that of current, state-of-the-art technologies, making it an optimal solution for the high mass, long distance architectures necessary for human spaceflight beyond cislunar space. NTP is the only currently viable option that delivers six astronauts to Mars within three months, reduces the total required number of launches, enables mission profiles with either shortterm or extended stays on the Martian surface, and offers abort scenarios at any point during transit (Sager, 1992;Durante and Bruno, 2010;Drake, 2013;Joyner, 2017). ...
... Having shown feasibility and affordability [1][2][3][4], the project focus has evolved as part of the STMD Space Nuclear Propulsion (SNP) Project within the Technology Demonstration Missions (TDM) Program. The current focus of SNP is in three major areas: ...
... Having shown feasibility and affordability [1,2,3,4], the project focus has evolved as part of the STMD Space Nuclear Propulsion (SNP) Project within the Technology Demonstration Missions (TDM) Program. The current focus of SNP is in three major areas: ...
... Having shown feasibility and affordability [1][2][3][4], the project focus has evolved as part of the STMD Space Nuclear Propulsion (SNP) Project within the Technology Demonstration Missions (TDM) Program. The current focus of SNP is in three major areas: ...
View Video Presentation: https://doi.org/10.2514/6.2021-3610.vid The long term, sustained exploration of Mars will require the focused efforts of a collective of national space agencies, public and private industry, and academia. The goal of this exploration will not be to just plant a flag and leave a boot print, but to continue beyond the first mission or first set of missions. Several key technologies will be required to enable both the upfront and continued exploration of Mars, one of these technologies being Nuclear Thermal Propulsion (NTP). This paper is the second in a series of studies analyzing a future Mars exploration campaign utilizing NTP as the primary propulsion option for both crew and cargo transfer. While the entire envisioned campaign includes uncrewed demonstration flights, crewed short-stay missions, and crewed long-stay missions, this paper examines the trajectory and architectural elements required for the first crewed mission to the Martian surface. Through a sequence of uncrewed and crewed missions, the Mars exploration campaign will prepare for the eventual permanent human presence on the planet’s surface. The first crewed mission to the Martian surface (third overall mission in the campaign) will be a short-stay mission utilizing an opposition class trajectory. With the proposed campaign structure, flying an opposition class trajectory for the first surface mission requires fewer logistics and surface assets to be developed and pre-deployed. This approach ensures a minimalist approach to the first surface mission, increasing reliability and safety, and spreading out the cost burden of an entire campaign. The elements required for this first mission would be the NTP Crew Mars Transport Vehicle (MTV), NTP Cargo MTV, Mars lander, Mars Ascent Vehicle (MAV), and surface habitation. The NTP Crew and Cargo MTVs would leverage design commonality across most subsystems, reducing cost and complexity. It is also envisioned that the MAV and Mars lander would have propulsion subsystem commonality. A combination of government and commercial launch vehicles are assumed for each element, utilizing the payload mass and volume capability of the Space Launch System (SLS) Block 2 along with the more frequent launch cadence of commercial vehicles. The crew transfer to Mars uses an opposition class trajectory with a total mission duration of 760 days, at the cost of higher DV and a larger transfer vehicle when compared to a conjunction class trajectory. The primary objectives of this mission would be to demonstrate the landing of large structures and short-term habitation on Mars, while also increasing experience with operating an NTP engine and human habitation in deep space.
... Having shown feasibility and affordability [1,2,3,4], the project focus has evolved as part of the STMD Space Nuclear Propulsion (SNP) Project within the Technology Demonstration Missions (TDM) Program. The current focus of SNP is in three major areas: ...
View Video Presentation: https://doi.org/10.2514/6.2021-3611.vid Since 2016, AR has been working with NASA and members of industry as part of the NASA Space Technology Mission Directorate. The initial goal of this project was to determine the feasibility and affordability of a low enriched uranium (LEU)-based nuclear thermal propulsion (NTP) engine with solid cost and schedule confidence. Having shown LEU NTP feasibility and affordability, the focus of the project has expanded to include not only NTP technology maturation, but also the identification of subsystem capability gaps and needs for another potential space nuclear propulsion technology: nuclear electric propulsion (NEP). As part of this activity, AR is examining potential in-space vehicle configurations leveraging a combination of NEP and chemical main propulsion to support a human Mars exploration campaign. This campaign consists of a series of missions, starting with uncrewed demonstration flights, crewed short-stay missions, and crewed long-stay missions preparing for the eventual permanent human presence on the surface of Mars. This paper will provide a brief overview of this envisioned campaign and detail the trajectory, the NEP/chem vehicle design concept, and the mission elements required for the first human Mars surface mission.
... This efficiency boost is particularly impactful for crewed missions beyond low-earth orbit (LEO) and NTP is a leading propulsion technology for a crewed mission to Mars. 2 NTP benefits include reduced travel times, additional abort capability and fewer heavy launch vehicles to enable the mission. 3 The fuel in the nuclear reactor core needs to withstand temperatures above 2500 K in hot hydrogen and thermal cycling. Ceramic-metallic fuels (cermets) can meet these requirements using a ceramic nuclear fuel (UO2 or UN) in a refractory metal matrix. ...
Nuclear thermal propulsion has both high thrust and high specific impulse and is a leading technology for a crewed Mars mission. Molybdenum cermets are an alternative to tungsten cermets that can reduce core mass and add ductility. The Mo matrix appears robust in a Mo-YSZ cermet after testing in hydrogen at 2500 K with thermal cycling. The subsurface Mo-YSZ interface also appears strong despite indications of debonding at the surface. Striations that appear parallel on the surface in the YSZ fuel surrogate extend below the surface.
... Since the pioneering work of Myers (2009), mathematical modeling of mission abort policies (MAPs) has received considerable research attentions in the past several years. These policies are designed to mitigate the risk of the system loss for lifecritical or safety-critical applications (Filene, 1974;Joosten, Drake, Weaver, & Soldner, 1991;Joyner, Kokan, Levack, Horton, & Widman, 2017). Specifically, in the case of certain deterioration condition defined by the MAP being met, the system abnegates its primary mission (PM) objective followed by a rescue procedure (RP) performed with the aim to survive the system (Levitin, Xing, & Huang, 2019). ...
For some critical applications, successfully accomplishing the mission or surviving the system through aborting the mission and performing a rescue procedure in the event of certain deterioration condition being satisfied are both pivotal. This has motivated considerable studies on mission abort policies (MAPs) to mitigate the risk of system loss in the past several years, especially for standby systems that use one or multiple standby sparing components to continue the mission when the online component fails, improving the mission success probability. The existing MAPs are mainly based on the number of failed online components ignoring the status of the standby components. This article makes contributions by modeling standby systems subject to MAPs that depend not only on the number of failed online components but also on the number of available standby components remaining. Further, dynamic MAPs considering another additional factor, the time elapsed from the mission beginning in the event of the mission abort decision making, are investigated. The solution methodology encompasses an event‐transition based numerical algorithm for evaluating the mission success probability and system survival probability of standby systems subject to the considered MAPs. Examples are provided to demonstrate the benefit of considering the state of standby components and elapsed operation time in obtaining more flexible MAPs.
... An NTP provides the most robust approach for future Mars and possibly round-trip lunar missions. 1,2,3,4,5 NASA and industry studies since the 1960's have indicated that nuclear thermal propulsion and fission power is a promising technology that has an established foundation that can be matured to flight status within the next ten years and that can be used for fast trip transfers for human missions to Mars or fast 24-48 hour taxis to the Moon. Many detailed mission and system studies have been performed that have shown the benefits in terms of payload, transfer time, and increased mission science via more power or reduced spacecraft mass. ...
Studies of Nuclear Thermal Propulsion (NTP) over the past several decades, and updated most recently with the examination of Low Enriched Uranium (LEU), have shown nuclear propulsion is an enabling technology to reach beyond this planet and establish permanent human outposts at Mars or rapidly travel to any other solar system body. The propulsion needed to propel human spacecraft needs high thrust to operate withinthe deep gravity well of a planet and provide high propulsive efficiency for rapid travel and reduced total spacecraft mass.
NTP can provide the thrust to move a spacecraft between orbits, can operate as a dual-mode system that provides power and propulsion capability, provides a strong architectural benefit to human and robotic exploration missions, and provides a path toward reusable in-space transportation systems. NTP provides smaller vehicle systems due to its specific impulse (ISP) being twice that of the best cryogenic liquid rocket propulsion and can thus provide reduced trip times for round-trip missions from Earth to Mars.
Aerojet Rocketdyne (AR) is working with NASA, other government agencies, and other industry partners to improve the design and reduce the cost of NTP engine systems. Current NTP designs focus on thrust sizes between 15,000-lbf (~67-kN) and 25,000-lbf (~111-kN). AR in 2019 has examined various reactor cores and enhancements to optimize LEU NTP designs. The enhancements improve the mission architecture robustness and provide more design margin for Mars vehicles across many mission opportunities, trip times, and mission types.
The LEU NTP design can offer mission architecture stages or elements that can be used for both Lunar and deep space exploration missions. The NTP designs enable packaging of various NTP stage designs (e.g., crew Mars vehicle stage elements, cargo stage derivatives, a deep space stage with payload, Lunar stage elements) on the NASA SLS Block 2 using the 8.4-meter fairing. The primary LEU core designs studied, for the above-mentioned missions, have relied on liquid hydrogen for the propellant and coolant and use Zirconium Hydride within a structural element as the neutron moderator. Several designs have been examined that use Beryllium Oxide with the fuel elements in the core to eliminate the structural elements.
The LEU NTP engine systems studied have typically been used only for the primary delta-V burns (e.g., earth escape, planetary capture, planetary escape, earth return capture). LEU NTP engine systems have also been examined using a the LEU reactor fuel element and moderator element approach to perform orbital maneuvering system (OMS) burns during the mission simply by permitting the reactor to keep operating at very low power levels during the entire mission.
This paper will discuss the various engine system and mission design trades performed in 2019 for Mars and lunar missions when using a single NTP or a cluster of NTP engines.
... The benefits of the technology are directly attributable to the increased specific impulse and high thrust of NTP (two times the best cryogenic propellant combinations, >15,000 lbf). Depending upon mission requirements, vehicle systems can be smaller, provide reduced transfer time, enable a return-to-Earth abort capability [2], or provide increased payload capability to their destinations. Taking advantage of these characteristics enables novel architectures to be developed for applications with respect to cis-lunar space, Mars missions, and deep space transport. ...
Since 2016, Aerojet Rocketdyne (AR) has been working with NASA and other industry partners to improve the design and increase the feasibility of Low Enriched Uranium (LEU) NTP engine systems that have clear programmatic and cost benefits over older Highly Enriched Uranium (HEU) designs. Part of this study effort is examining potential value-added deep-space architectures where NTP can be incorporated beyond Mars exploration. The baseline Mars Point-of-Departure (POD) crewed vehicle design includes a core stage with three LEU NTP engines, each producing 25,000 lbf of thrust, and three common inline stages. A NTP cargo vehicle was analyzed for the delivery of Mars surface assets. It has a single LEU NTP engine and has common systems with the Mars POD crew vehicle. By leveraging the design of this cargo vehicle, a lunar architecture that incorporates NTP can be developed while proving propulsion technology that enables the future exploration of Mars and the Outer Solar System.
Several lunar Design Reference Missions (DRMs) were developed to cover a wide range of payloads, destinations, and trajectory options while taking advantage of the benefits that NTP provides. For each DRM, NASA’s Space Launch System (SLS) was used to position the NTP transfer vehicle into Earth orbit, with commercial launch vehicles available for propellant resupply. In DRMs 1 and 2, the NTP transfer vehicle delivers payload directly into Low Lunar Orbit (LLO) and Near Rectilinear Halo Orbit (NRHO), respectively, and returns to the launch vehicle drop-off orbit ready for refueling and a new payload. For each trajectory, the NTP transfer vehicle utilizes either Weak Stability Boundary (WSB) transfers or lunar direct transfers based on the requirements of the mission and payload mass.
For DRM 3, the NTP transfer vehicle provides transport between NRHO and LLO, acting as an in-space stage for either payload/propellant delivery or lunar lander/crew transport. NTP’s high performance allows for unique transfer options while maintaining the ability to perform multiple missions without refueling. Finally, DRM 4 utilizes a free return transfer option between the Moon and Earth orbit for either civilian tourism or payload drop-off, where the payload performs the Lunar Orbit Insertion (LOI) via a bi-elliptic transfer. Upon return to Earth orbit, the NTP transfer vehicle awaits a new payload and, if needed, is refueled. The Design Reference Missions detailed in this paper show how NTP can be incorporated into a lunar architecture for cargo/crew delivery to cis-lunar space, as well as transport between NRHO and LLO.
View Video Presentation: https://doi.org/10.2514/6.2022-4374.vid Throughout the entire history of human spaceflight, astronauts have never been more than a few days’ journey from Earth. For missions to the International Space Station, or even to the Moon, aborting the mission and returning home is a relatively straightforward operation in the event of a life-threatening contingency. But on the journey to Mars, a mission abort is complicated by a single factor: the sheer distance to Mars that will take time to cover regardless of how powerful the propulsion system is. In a significant departure from the entire human spaceflight experience to date, a Mars mission abort will not mean an immediate return to Earth due to the heliocentric nature of the transit. After escaping Earth’s gravity and entering heliocentric space, aborting a mission to Mars will be measured in months not days even if the abort is initiated relatively early in the mission. This represents a fundamental shift in the way mission aborts are utilized to reduce mission risks and this change must be part of the narrative when discussing reliability, crew risk, contingency planning, and mission operations. Abort capability will vary between transportation propulsion systems depending on final design, operational implementation, and actual mission parameters such as departure date, but the difference is relatively insignificant given that “mission abort” in the heliocentric frame of reference is a wildly different paradigm than that historically used for Earth-centric missions which rely on abort as a risk mitigation strategy. The primary objective of this paper is to provide a clearer definition of human Mars mission aborts from a celestial physics and orbital mechanics perspective in the context of different propulsion systems.
Modeling and optimizing mission abort policies have received intensive research attention in the past decade. While most of the existing models assumed single-attempt missions, few of them considered multi-attempt missions (i.e., a system component may re-attempt the mission after being rescued successfully and properly maintained). However, the existing multi-attempt studies assumed the same abort policy (AP) for all the system components. This paper extends the state of the art by putting forward a new shock-based AP model where a subset of available components during each attempt is identified as kamikaze components that operate with a riskier AP than the rest of the available components, and the number of kamikaze components and APs for both groups of components (kamikaze and non-kamikaze) may change from attempt to attempt. The AP optimization problem is solved using the genetic algorithm to minimize the total expected cost of losses. The problem is further generalized to incorporate the total number of components and the maximum number of mission attempts as additional decision variables. The proposed AP and solutions to both optimization problems are illustrated through a detailed case study of a multi-UAV system undergoing random shocks during both the primary surveillance mission and the rescue procedure.
Though intensive research efforts have been devoted to the study of mission aborting rules (MARs) for diverse types of systems, no previous work has discussed this problem for systems with product storage. This paper contributes by modeling and optimizing MARs for multistate systems with product storage, which may accumulate surplus product when the system performance surpasses the required demand and compensate product deficiency otherwise. The storage has limited capacity, uploading performance, and downloading performance. The system state may deteriorate due to random shocks, which have different severity and inter-arrival time distributions during primary mission (PM) and rescue procedure (RP). We put forward a numerical algorithm to evaluate the mission success probability (MSP) and damage avoidance probability (DAP) of the considered system with specified PM and RP demands and durations. Based on the MSP and DAP evaluation, we make a further contribution by formulating and solving a constrained optimization problem that finds the optimal MAR maximizing the MSP while meeting a desired level of DAP. Using a power system example, we also investigate influences of storage capacity, storage uploading/downloading performance constraints, and desired DAP level on the system performance metrics and on the MAR optimization solutions.
Mission aborting has recently attracted great attention, where mission abort rules (MARs) have been modeled and optimized for different types of technological systems aiming to effectively mitigate the risk of system losses. However, none of the existing works have considered systems with multiple work-sharing units. This article makes advancements in the state of the art by modeling condition-based MARs for nonrepairable work-sharing systems that must perform a specified amount of work during the primary mission (PM). The MAR considered presumes aborting the PM to prevent considerable damage when the number of available units reduces to a certain number
while the amount of work accomplished in the PM is less than
(
). After the PM abortion, a rescue procedure (RP) is executed by the remaining units to survive the system. Dynamic operating conditions during PM and RP are considered. A probabilistic model-based numerical algorithm is proposed to evaluate several performance metrics, including mission success probability, RP success probability, damage avoidance probability, and expected cost of losses. The MAR optimization problem for minimizing the expected cost of losses is formulated and solved using the genetic algorithm. An example of a chemical reactor system is provided to demonstrate the proposed algorithm as well as the benefit of MARs optimized in comparison to the “no abort” policy and the most conservative abort policy (the PM aborts upon the failure of any unit). Effects of several model parameters on the system performance metrics and optimization solutions are also examined through examples.
This paper models a parallel system with multiple units performing and sharing a specified amount of work required for the system's primary mission (PM). The system operating environment is exposed to random shocks that may deteriorate or fail the units. To balance the risk of losses/damages caused by the mission failure and the amount of accomplished work, a dynamic task distribution policy (TDP) is proposed that, after the occurrence of each shock, determines the distribution of available units between performing the PM work and performing a damage reduction procedure (DRP). While this work belongs to the line of recent research on mission aborting, it makes advancement in the state of the art by considering the partial abort through the PM and DRP work redistribution among the available units (different from the PM full abort of existent models) and effects of external shocks for work sharing systems. The specific contributions made in this work include a discrete numerical algorithm for evaluating the PM success probability, expected fraction of the completed PM work, and expected total cost of losses (ECL) for the considered parallel system subject to the dynamic TDP. The TDP optimization problem is also formulated and solved, leading to the minimized ECL. Application of the proposed model and effects of several model parameters are demonstrated through a liquid supply system in the chemical reactor and through detailed analyses and optimizations of a multi-processor data processing system.
Mission aborting aims to survive a system and avert catastrophic damages caused by the loss of the system. However, the effectiveness of mission aborting depends heavily on the abort policy adopted. Thus, it is crucial to model the abort policy in the mission performance evaluation and further optimize the policy maximizing or minimizing the mission performance metric of interest. Rich research efforts have recently been expended in studying the abort policy, and all the existing models have assumed the full mission abort in the case of a pre-determined condition being met. This paper makes novel contributions by modeling and optimizing a partial abort policy (PAP) for systems with multiple work sharing units. With the goal of balancing the amount of completed work and the risk of damages, the PAP proposed determines the number of units that should continue the primary mission (PM) whereas the rest of the operating units abort the PM and start the execution of a rescue procedure (RP). The traditional full abort policies appear as special cases of the proposed PAP model. A numerical algorithm based on probabilistic models is suggested for mission performance evaluation while considering dynamic operation conditions of units during PM and RP. Mission performance metrics evaluated include mission completion probability, damage avoidance probability, expected completed work fraction, and expected losses. The optimal PAP problem is further formulated and solved using the genetic algorithm with the objective of minimizing the expected losses. Influences of multiple model parameters on the mission performance metrics and optimization solutions are examined through an example of a multi-processor server system.
This study examines the impact of the launch capabilities provided by the planned Space Launch System (SLS) Block 2 vehicle on the solar electric propulsion (SEP) power levels required for large cargo prepositioning missions to Mars orbit. Our analysis shows that the additional launch capability planned for the SLS vehicle enables a dramatic reduction in the required SEP vehicle power level while maintaining reasonable trip times for large cargo pre-positioning (20-40mt). For near-term missions, we show that 50kW SEP modules can be used either singly or in combination, enabling much lower cost human exploration missions. This is a much lower power level, and thus much lower cost, than shown by our previous studies. 50kW SEP modules can also be used for cis-lunar missions and large science missions such as the Asteroid Redirect mission with smaller launchers such as the SLS Block 1, Falcon 9 and 9H, and EELV launchers, ensuring that the SEP modules have multiple applications. We show that this modular approach enables a significant reduction in total development and production costs for high power SEP and provides for a more gradual phasing of system development to fit within available annual budgets.
As the nation embarks on a new and bold journey to Mars, significant work is being done to determine what that mission and those architectural elements will look like. The Evolvable Mars Campaign, or EMC, is being evaluated as a potential approach to getting humans to Mars. Built on the premise of leveraging current technology investments and maximizing element commonality to reduce cost and development schedule, the EMC transportation architecture is focused on developing the elements required to move crew and equipment to Mars as efficiently and effectively as possible both from a performance and a programmatic standpoint. Over the last 18 months the team has been evaluating potential options for those transportation elements. One of the key aspects of the EMC is leveraging investments being made today in missions like the Asteroid Redirect Mission (ARM) mission using derived versions of the Solar Electric Propulsion (SEP) propulsion systems and coupling them with other chemical propulsion elements that maximize commonality across the architecture between both transportation and Mars operations elements. This paper outlines the broad trade space being evaluated including the different technologies being assessed for transportation elements and how those elements are assembled into an architecture. Impacts to potential operational scenarios at Mars are also investigated. Trades are being made on the size and power level of the SEP vehicle for delivering cargo as well as the size of the chemical propulsion systems and various mission aspects including In-space assembly and sequencing. Maximizing payload delivery to Mars with the SEP vehicle will better support the operational scenarios at Mars by enabling the delivery of landers and habitation elements that are appropriately sized for the mission. The purpose of this investigation is not to find the solution but rather a suite of solutions with potential application to the challenge of sending cargo and crew to Mars. The goal is that, by building an architecture intelligently with all aspects considered, the sustainable Mars program wisely invests limited resources enabling a long-term human Mars exploration program. Nomenclature C3 = Escape Energy (km 2 /s 2) R = Distance from the Sun V = Propulsive Delta Velocity (m/s)
Trajectory options for conjunction-class human Mars missions are examined, including crewed Earth-Mars trajectories with the option for abort to Earth, with the intent of serving as a resource for mission designers. An analysis of the impact of Earth and Mars entry velocities on aeroassist systems is included, and constraints are suggested for interplanetary trajectories based upon aeroassist system capabilities.
Mars trajectory design options were examined that would accommodate a premature termination of a nominal manned opposition class mission for opportunities between 2010 and 2025. A successful abort must provide a safe return to Earth in the shortest possible time consistent with mission constraints. In this study, aborts that provided a minimum increase in the initial vehicle mass in low Earth orbit (IMLEO) were identified by locating direct transfer nominal missions and nominal missions including an outbound or inbound Venus swing-by that minimized IMLEO. The ease with which these missions could be aborted while meeting propulsion and time constraints was investigated by examining free return (unpowered) and powered aborts. Further reductions in trip time were made to some aborts by the addition or removal of an inbound Venus swing-by. The results show that, although few free return aborts met the specified constraints, 85% of each nominal mission could be aborted as a powered abort without an increase in propellant. Also, in many cases, the addition or removal of a Venus swing-by increased the number of abort opportunities or decreased the total trip time during an abort.
Determining Mars Mission NTP Thrust Size and Architecture Impact for Mission Options
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