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Impacts of In Situ Alternative Propellant on Nuclear Thermal Propulsion Mars Vehicle Architectures
Abstract
Recently, NASA has pushed for returning humans to the moon sustainably with In situ resource utilization being the central focus. In its permanently shadowed regions, the moon has an abundance of water and ammonia, which are two potential alternative propellants to hydrogen in nuclear thermal propulsion (NTP) engines. Using Aerojet Rocketdyne Mars mission architectures and University of Alabama in Huntsville NTP engine models, this research analyzed the impacts of using water or ammonia as propellants on mission architectures. For a human mission to Mars originating in the lunar distant retrograde parking orbit, when comparing the baseline hydrogen vehicles to vehicles using water or ammonia, the effects of increased propellant density by an order of magnitude despite a 62% decrease in specific impulse will still require an average of 53% of the launches for vehicle assembly, 59% of the dry mass, 25% of the propellant volume for the conjunction-class mission, and 50% of the propellant volume for the opposition-class mission when comparing against the baseline hydrogen vehicle. Due to the oxidative nature of water, which results in 29% of the engine life compared with that of hydrogen and ammonia, ammonia outperforms water in terms of the total number of missions per engine block by an average factor of 6.8. Because ammonia’s specific impulse is 40.6% that of hydrogen, the propellant mass is increased by a factor of 2.9 and longer burn times result in an average of a 10% decrease in the total number of missions that the hydrogen system can achieve. Since the changes to the vehicle architecture are more extreme and in favor of ammonia than the total mission capability, for reusable vehicles using NTP technology ferrying humans to Mars while utilizing in situ propellants from the moon, this research recommends ammonia as the propellant of choice.
... Nevertheless, in the years, other propellants have been proposed for the NTP, to overcome some drawbacks presented by the use of the hydrogen, as its very low density and the need to be stored in cryogenic conditions. The propellant most proposed in the literature [3,21,22] are reported in Table 6, together with their main properties and the specific impulse estimated at 3000 K. ...
... These particles embedded in a ZrC matrix gives the selected CerCer material for this preliminary trade-off. The low uranium content assures a melting temperature in the order of 3800 K and the density will be relatively low compared with other types of fuels ranging from 7-9 g/cm 3 . ...
... The UO2 Cermet fuel of choice is the 40% vol W-60% vol UO2 with 5 mol.% Gd2O3 and W coating, presenting a melting temperature over 3100 K and a density of about 14 g/cm 3 . ...
From the late 1950s to date, nuclear thermal propulsion has the purpose of making possible missions impossible to accomplish with chemical propulsion. The use of hydrogen as a propellant and a fuel resistant to very high temperatures guarantees exceptional propulsive performance. Since the NERVA project, a propulsion system with such capabilities has proven to be the most suitable to carry out missions to Mars. Even today, the literature on nuclear thermal propulsion systems appears to be full of Mars mission studies. Little attention is given to other missions potentially made accessible by this type of propulsion. The purpose of this work is to investigate different mission scenarios that could be made accessible by nuclear propulsion and to identify the most efficient configurations of the propulsion system for each scenario. Some mission scenarios are presented, covering both interplanetary and earth orbit missions. The specific requirements that each mission imposes on the propulsion system are determined. Particular attention is given to the analysis of the safety requirements, which constitute the most stringent constraints to be respected for the use of nuclear reactors in space. A literature review identifies the most promising fuels and propellants, with a focus on the recently developed high temperature fuels. A preliminary design determines the mass, size and materials of the components of the propulsion systems satisfying the imposed requirements for the diverse missions. All the proposed configurations have associated different combinations of fuels and propellants among the one chosen from the previous literature reviews. For each mission, a trade-off is made based on the performance offered by the various propulsion systems proposed. Particular attention is given to the parameter of the system specific impulse, which weighs the specific impulse on the mass of the propulsion system. This parameter is very important in systems using nuclear propulsion, where the mass of the engine occupies a large part of the total mass of the system. From this preliminary trade-off, the nuclear thermal propulsion system configurations worth to be studied in detail and optimized emerge for each identified mission. Finally, the possible benefits in terms of system specific impulse derived by the use of the fission reactor also for power generation purposes are discussed.
... There are also other methods of hydrogen and oxygen separation, such as the sulfur-iodine cycle [22]; however, the consideration of these methods is deferred to future research. Previous works [23,24] compared the mission performance metrics of using alternative propellant NTP (A-NTP), combustion, and liquid-oxygen augmented nuclear thermal rocket (LANTR) engines against NASA's baseline hydrogen-based NTP (H-NTP) engines for both Mars conjunction and opposition class missions. The results showed that LANTR engines yielded the best overall mission performance when propellant availability for retanking the vehicles between each mission was assumed while noting that both H-NTP and A-NTP engines are important stepping stones for technology demonstration, which will lead to the development of secondgeneration NTP engines such as the LANTR. ...
... The stages are then transferred to LDRO for aggregation. The conjunction class MTV consists of a core stage with three H-NTP engines and some propellant, three inline stages with propellant and maneuvering systems, and a habitat all stacked together like a single-stage rocket [11,23,25]. ...
... (3)] to replace the propellant mass in Eq. (1). The full derivation was provided in a previous work [23]. Furthermore, to capture the mass of the propulsion system and the energy used for a single mission in one term, the energy was divided by the mass of the propulsion system to yield the specific energy of the system E sp [16][17][18][19]28]. ...
Recently, NASA has pushed for returning humans to the moon sustainably with in situ resource utilization as the central focus. The moon has an abundance of water that is proposed to be electrolyzed into hydrogen and oxygen to be used as propellant. Other volatiles such as ammonia, carbon dioxide, and methane are also present. A mission architecture for a lunar ascent/descent vehicle (LADV) from the Polytechnic University of Turin and nuclear thermal propulsion (NTP) engine models from the University of Alabama in Huntsville were used to compare in-situ-derived propellants for a LADV. This study considered a LADV originating from the lunar surface, delivering a payload in the lunar distant retrograde orbit, and returning to the lunar surface for retanking. This research analyzed the impacts on this mission of using hydrogen NTP, water/ammonia NTP, liquid-oxygen augmented nuclear thermal rocket, and Aeon 1 methane–oxygen engines using the selected architecture and tools. The results were compared to the reference hydrogen–oxygen RL10 engine. The propulsion system comparison analysis showed that combustion engines will offer better overall performance than NTP-based engines due to a 50% decrease in propellant volume, a 20% decrease in dry mass, and a lower propellant mass than the water and ammonia NTP systems. Both the hydrogen–oxygen and methane–oxygen propulsion systems will have similar propellant masses when compared to other systems. This is due to the order of magnitude higher mass of the NTP engines, with the highest mass contribution coming from the reactor. However, both water and ammonia alternative propellant NTP engines can still be viable candidates for the usage of these minimally processed propellants to satisfy this mission.
... (3)] to replace the propellant mass in Eq. (1). The full derivation was provided in previous work [11]. Furthermore, to capture the mass of the propulsion system and the energy used for a single mission in one term, the energy was divided by the mass of the propulsion system to yield the specific energy of the system E sp [10][11][12][13][14][15]. ...
... The full derivation was provided in previous work [11]. Furthermore, to capture the mass of the propulsion system and the energy used for a single mission in one term, the energy was divided by the mass of the propulsion system to yield the specific energy of the system E sp [10][11][12][13][14][15]. ...
... This study focuses on vehicles using RL10 for LH2 liquid-oxygen (LOX) propellant [23], Relativity Space's Aeon-1 for LCH4 LOX propellant [24], and LOX Augmented Nuclear Thermal Rocket (LANTR) [25] engines, which are compared with the reference vehicles utilizing H-NTP engines. Use of water and ammonia directly in NTP engines has been previously considered [11]. Fig. 4 Opposition-class pure-hydrogen MTV CONOPS (adapted from AR's Mars opposition vehicle architecture [22]). ...
Recently, NASA has pushed for returning humans to the moon, with in-situ resource utilization being the key capability to provide sustainability. One of the potential future developments could be a propellant depot in lunar distant retrograde orbit. Using Aerojet Rocketdyne Mars mission architectures and University of Alabama in Huntsville Nuclear Thermal Propulsion (NTP) engine models, this research analyzed the impacts of using chemical [Formula: see text] and [Formula: see text] engines as well as the liquid-oxygen (LOX) Augmented Nuclear Thermal Rocket (LANTR) engines for these missions and compared their performances to the reference hydrogen-based NTP (H-NTP) engines all the while assuming a propellant depot at lunar distant retrograde orbit. For a human mission to Mars originating in the lunar distant retrograde parking orbit, the LANTR engines will offer better overall performance than H-NTP engines with a predicted 55.6% decrease in propellant volume, 39% decrease in vehicle dry mass, and 50% decrease in the number of aggregation launches. This is due to LANTR’s 22% higher specific impulse than conventional [Formula: see text] chemical propulsion systems, three times higher density than pure hydrogen, and 440% higher thrust than the baseline H-NTP engines. However, these benefits come at the cost of the propellant mass, which is 32.4% higher for the conjunction class mission and 106.7% higher for the opposition class mission than the baseline H-NTP system.
... An overview of NTPs for Mars travel vehicles is provided in [16], while the authors in [17,18] compare the performance of different NTP engines for lunar and Mars applications. The work in [19] demonstrated that while using NTP (specific impulse 900 s) and NEP (specific impulse of 6000 s), there is significant LEO mass reduction compared to chemical propulsion (specific impulse of 440 s). ...
... An overview of NTPs for Mars travel vehicles is provided in [16], while the authors in [17,18] compare the performance of different NTP engines for lunar and Mars applications. The work in [19] demonstrated that while using NTP (specific impulse 900 s) and NEP (specific impulse of 6000 s), there is significant LEO mass reduction compared to chemical propulsion (specific impulse of 440 s). A multi-MW NEP is discussed in [20] with a focus on mass reduction, while the work in [21][22][23] studied the possibility of NEPs combined with a chemical propulsion stage for human Mars-mission spacecraft. ...
This paper proposes a megawatt (MW)-scale high-voltage (HV) electrical power-conversion element for large-spacecraft electric propulsion (EP) systems. The proposed scheme is intended for long-term and crewed missions, and it is driven by a nuclear electric propulsion (NEP) that acts as a heat source. The scheme includes (i) A two-rotor generator (TRG), (ii) A rectification stage, and (iii) An isolated dual output DC-DC (iDC2) converter. The TRG is a high-reliability electric machine with two rotors, a permanent magnet rotor (PMR), and a wound field rotor (WFR). The PMR has a fixed flux and hence back-EMF, while the back-EMF due to the WFR is controlled by injecting a direct current (DC) into the WFR winding. The total TRG output voltage, which is the sum of voltages due to the PMR and WFR, is controlled over a prescribed region of spacecraft operation. The output of the TRG is rectified and connected to the input of the iDC2 converter. The iDC2 converter uses a three-winding transformer, where the primary winding is fed from the rectified output of TRG, the secondary winding processes the propulsion power to an electric thruster via a high-voltage DC (HVDC) link and a tertiary winding that is connected to the spacecraft’s low-voltage DC (LVDC) power system. Three controllers are proposed for the system: an HVDC voltage controller, an HVDC current controller that controls the voltage and current processed to the thruster, and an LVDC controller that adjusts the current to the LVDC system. Detailed analytical models for the TRG, iDC2 converter, and controllers are developed and verified via simulations under different conditions. The analytical studies are further validated via results from a laboratory prototype.
... This chapter discusses the results of NTP engine trades for robotic missions to the outer planets. As discussed in the DRM trade tree, the missions to propellants such a ammonia due to their increased density and ease of long-term storage [110]. There have been limited peer reviewed studies on the long term inspace storage of LH2 for duration of five or more years, therefore, the architectures for Ice giant missions is restricted to the expendable mission mode. ...
... The point design studies and engine trades were limited to the use of LH 2 as a propellant for the NTP system. Although studies have shown benefits for propellants such as ammonia due to their increased density and ease of long-term storage [36]. Some of the leads on future work that can be used to expand the understanding of the NTP system for robotic missions will be to perform engine trades using alternative propellants and architecture evaluation for missions to ocean worlds with high ΔV requirements during planetary orbit insertion and use of the bimodal NTP concept. ...
Nuclear thermal propulsion (NTP) has emerged as a promising technology for enhancing the capabilities of robotic missions in space exploration. This paper investigates and analyzes the tradeoffs and sensitivity associated with utilizing NTP systems for robotic missions to the outer planets. The engine trade results for Jupiter and Neptune rendezvous missions using expendable configuration have demonstrated that the thrust range of 12.5–15 k-lbf (55.6–66.7 kN) can enable new frontiers and flagship-class missions with minimum initial mass in low Earth orbit. The specific impulse sensitivity analysis has shown that using a 13 k-lbf (57.8 kN) engine with an [Formula: see text] as low as 850 s can enable flagship-class missions to the gas giants in a direct transfer trajectory.
... The Complex Systems Integration Laboratory (CSIL) at the University of Alabama in Huntsville (UAH) has carried out multiple point design studies on Nuclear Thermal Propulsion (NTP) for missions in cislunar space, Mars, and for the outer planet exploration [27][28][29][30][31][32]. These studies utilized the Spacecraft Integration System Model (SISM), which encompasses systems engineering architecture models and domain-specific analysis models [33][34][35][36][37][38] . ...
The exploration of outer planets and icy moons demands advanced propulsion and power solutions to overcome the limitations of chemical and solar-electric systems. This paper highlights the transformative potential of Nuclear Thermal Propulsion (NTP) and nuclear power systems for deep-space missions. NTP systems, with their high thrust and specific impulse, enable reduced trip times and increased payload capacity, facilitating flagship-class missions to distant worlds. Nuclear power systems provide sustained energy for high-power instruments, such as ice-penetrating radars, LIDARs, and mass spectrometers, critical for investigating subsurface oceans and detecting organic compounds. Additionally, they enable high-rate data transmission over vast distances, ensuring maximum scientific return from robotic missions to the outer solar system.
... In this mission, NEP/NTP is responsible for low-thrust propulsion for most of the mission, while chemical propulsion, such as liquid oxygen/methane-based systems, provides high-thrust maneuvers near Earth/Mars [26]. An overview of NTPs for Mars travel vehicles is provided in [27], while [28]- [29] compare the performance of different NTP engines for lunar and Mars applications. The work in [30] demonstrates that utilizing NTP (specific impulse 900 s) and NEP (specific impulse 6000 s) reduces the mass in LEO compared to chemical propulsion (specific impulse 440 s). ...
This paper proposes a generalized control for multiphase multilevel neutral-point-clamped (MPML-NPC) converter in a large spacecraft electric propulsion (EP) system. The spacecraft EP scheme considered in this paper has been proposed by NASA for long-term crewed missions and includes a heat source that acts as a prime mover to a fixed speed permanent magnet generator (PMG) whose output is rectified via a MPML-NPC converter before connecting to a high-power electric thruster. The proposed generalized control is based on space vector modulation (SVM) developed for MPML-NPC converters with any number of phases and levels. The paper will discuss mathematical modelling of multiphase systems and derive the equations that govern the proposed control. The generalized control generates optimized switching commands based on a space vector diagram (SVD). The analytical SVM models are verified through simulations for 3-phase, 6-phase, and 9-phase MPML-NPC converters. Further, the simulations are validated through test results from a scale-down laboratory prototype MPML-NPC converter.
View Video Presentation: https://doi.org/10.2514/6.2022-4354.vid In the evolving landscape of contemporary space exploration, mission architectures must target programmatic objectives and constraints alongside mission performance requirements such that a portfolio of systems can be selected for a space campaign. Development and acquisition remain a gap in the programmatic assessment of space systems and campaigns. The operational gap for systems development and acquisition can create scheduling slippages that render entire campaigns infeasible or unviable due to the high temporal sensitivity of space exploration. A conceptual phase analysis for space systems development and acquisition is key to adding necessary programmatic consideration upfront for architecting space campaigns. This paper presents a mixed integer programming (MIP) formulation for modeling space systems development and acquisition, based on a strategic planning perspective of space campaigns. The problem is decomposed into stakeholders, missions, and capabilities as representation of a multi-level analysis that is architecture-agnostic. Stakeholders provide a priority vector of interests, missions fulfill certain interests and have capabilities as co-requisites for execution, and capabilities require resource allocation to be developed and sustained for usage. The formulation is a multi-period precedence-constrained knapsack problem (MPPCKP), where selection of missions and dependent capabilities must maximize value for a stakeholder’s interest fulfillment, subject to periodic budgetary constraints. Optimization results yield a portfolio of mission and capabilities, along with development timelines and budget allocations which can be represented via a sand chart. Demonstration of this formulation is shown using a sample problem and slight variations of based on current cislunar studies. Results showcase the speed and generality of the formulation with respect to multiple scenarios; this model lends itself to many conceptual phase space campaign analyses—of which one is its inclusion in a cislunar landscape tabletop forecasting environment.
View Video Presentation: https://doi.org/10.2514/6.2021-3612.vid This research illustrates the impact of different technology choices and their associated performance on overall sizing for a hybrid Nuclear Electric Propulsion (NEP) and chemical crewed Mars transportation vehicle. A set of models for the trajectory and vehicle mass, power system specific mass (alpha), and power conversion performance enable the analysis. Vehicle mass required to close the mission is calculated as a function of alpha and specific impulse (Isp) for multiple choices of high-level assumptions including system alpha vs. power scaling and electric thruster type. System alpha is built up and compared for multiple cases of technology assumptions. A Brayton cycle model is used to illustrate the alpha sensitivities to assumptions in cycle design such as peak operating temperature and component performance. The analysis illustrates how technology performance sensitivities affect the alpha and thus vehicle sizing for a range of mission and technology assumptions.
The future of human exploration missions to Mars is dependent on solutions to the technology challenges being worked on by the National Aeronautics and Space Administration (NASA) and industry. One of the key architecture technologies involves propulsion that can transport the human crew from Earth orbit to other planets and back to Earth with the lowest risk to crew and the mission. Nuclear thermal propulsion (NTP) is a proven technology that provides the performance required to enable benefits in greater payload mass, shorter transit time, wider launch windows, and rapid mission aborts due to its high specific impulse and high thrust.
Aerojet Rocketdyne (AR) has stayed engaged for several decades in working NTP engine systems and has worked with NASA recently to perform an extensive study on using low-enriched uranium NTP engine systems for a Mars campaign involving crewed missions from the 2030s through the 2050s. Aerojet Rocketdyne has used a consistent set of NASA ground rules and they are constantly updated as NASA adjusts its sights on obtaining a path to Mars, now via the Lunar Operations Platform-Gateway. Building on NASA’s work, AR has assessed NTP as the high-thrust propulsion option to transport the crew by looking at how it can provide more mission capability than chemical or other propulsion systems.
The impacts of the NTP engine system on the Mars transfer vehicle configuration have been assessed via several trade studies since 2016, including thrust size, number of engine systems, liquid hydrogen stage size, reaction control system sizing, propellant losses, NASA Space Launch System (SLS) payload fairing size impact, and aggregation orbit.
An AR study activity in 2018 included examining NTP stages derived from Mars crew mission elements to deliver extremely large cargo via multiple launches or directly off the NASA SLS. This paper provides an update on the results of the ongoing engine system and mission trade studies.
DOE Position Paper on the potential role of ammonia in a hydrogen economy. Focused on vehicular applications
Aside ² from Earth, the inner solar system is like a vast desert where water and other volatiles are scarce. An old saying is, “In the desert, gold is useless and water is priceless.” While water is common on Earth, it is of very high value in space. Science missions to the Moon have provided direct evidence that regions near the lunar poles, which are permanently in shadow, contain substantial concentrations of water ice. On the lunar surface, water itself is critical for human consumption and radiation shielding, but water can also be decomposed into hydrogen and oxygen via electrolysis. The oxygen thus produced can be used for life support, and hydrogen and oxygen can be combusted for rocket propulsion. Due to the Moon's shallow gravity well, its water-derived products can be exported to fuel entirely new economic opportunities in space. This paper is the result of an examination by industry, government, and academic experts of the approach, challenges, and payoffs of a private business that harvests and processes lunar ice as the foundation of a lunar, cislunar (between the Earth and the Moon), and Earth-orbiting economy. A key assumption of this analysis is that all work—construction, operation, transport, maintenance and repair—is done by robotic systems. No human presence is required. Obtaining more data on conditions within the shadowed regions is vital to the design of a lunar ice processing plant. How much water is actually present, and at what percentage in the lunar regolith? How firm or soft are the crater bottoms, and how will that affect surface transportation? How deep is the ice resource, and in what state is it deposited amongst the regolith? These and other questions must be answered by precursor prospecting and science missions. A wide range of potential customers for the hydrogen and oxygen products has been identified. They can be used to fuel reusable landers going back and forth between the lunar surface and lunar orbit. They can make travel to Mars less expensive if the interplanetary vehicle can be refueled in cislunar space prior to departure. Operations closer to Earth can also benefit from this new, inexpensive source of propellant. Refueling in Low Earth Orbit can greatly improve the size, type, and cost of missions to Geosynchronous Earth Orbit and beyond. This study has identified a near term annual demand of 450 metric tons of lunar derived propellant equating to 2450 metric tons of processed lunar water generating 4 billion, about the cost of a luxury hotel in Las Vegas. With this investment however, a scalable market can be accessed. As refueling decreases in-space transportation costs, entirely new business and exploration opportunities will emerge with potential to vastly benefit the economies of Earth. Even with the early customers identified within this study, it has been determined that this could be a profitable investment with excellent growth opportunities. The United States Government has critical roles to play in the development of this commercial capability as well. Government science/prospecting and communications missions to the Moon can be very helpful in both the development and operational phases of the business. Government laboratories can contribute some of their technologies and help facilitate integrated systems tests of a terrestrial pilot plant. Government must also work to fill the gaps in international law regarding property rights on celestial bodies such as the Moon. In addition, between Earth orbit, Moon, and Mars missions, government could be an important anchor customer for the resource, stimulating the private sector into action with proposed demands and price points while improving its mission costs and capabilities. This study demonstrates both the technical and economic feasibility of establishing a commercial lunar propellant production capability. It provides recommendations to interested government and private organizations and defines a path to implementation; and explains that by doing so the United States will fuel a new age of economic expansion, sustained space exploration, settlement, and American leadership in space.
Performance up to 30 kW discharge power of a 100-kW class, 3-channel nested Hall thruster operating on xenon propellant is presented. Each of the seven combinations of thruster channels were run at 75% of the nominal thruster current density and 300 V anode potential. Discharge power ranged from 4 kW for the inner channel on its own to 30 kW for all three channels operating together. Thrust was measured using a new inverted-pendulum, null-type thrust stand designed to safely support over 250 kg of thruster mass while accurately resolving thrust. With all 3 channels running at 30 kW total discharge power, the anode specific impulse was 1840 s, anode efficiency was 45.0%, and thrust was 1518 mN. Anode specific impulse ranged from 1200 s to 2300 s across operating conditions and anode efficiency ranged from 21% to 67%. A discussion of the effect of chamber backpressure on performance measurements is included. Comparison of X3 performance is made to other state-of-the-art Hall thrusters operating at similar fractions of current density.
Recently, NASA’s Artemis Program has considered sustainably returning humans to the moon with in situ resource utilization as the central focus and considers nuclear thermal propulsion (NTP) to be a viable propulsion system candidate for deep space missions. This program also considers the moon to function as a staging area for deep space missions due to its strategic orbit around Earth. The Lunar Crater Observation and Sensing Satellite discovered the abundancy of water, ammonia, and other volatiles in permanently shadowed regions on the moon. Because NTP can theoretically use any fluid as a propellant, water and ammonia are two potential alternative propellants to hydrogen that do not require postprocessing to be used directly in situ. Alternative propellant NTP engine models were developed to examine the performance of water and ammonia. It was found that bleed cycle architectures are infeasible with current turbine materials, but expander cycles have acceptable turbine inlet temperatures with marginal performance differences from bleed cycles. At reactor power levels comparable to hydrogen NTP engines, water doubles the thrust while ammonia increases it by 70%, although at a much lower specific impulse. The largest contributor to the life of the engine was the maximum temperature of the SiC cladding for water and the maximum temperature of the fuel for ammonia where higher values resulted in shorter engine life.
Seeding hydrogen is the process of adding a heavy noble gas (the seed) such as argon, krypton, or xenon to the hydrogen propellant. This adds another degree of freedom to the design of a nuclear thermal propulsion (NTP) engine. This is done to reduce pressure losses, improve convective heat transfer, and densify the propellant at the expense of specific impulse and wetted vehicle mass. A numerical study was conducted which predicted and examined the effects of varying the seed concentrations within the hydrogen propellant in the steady state operation of an NTP engine. The current study examines the transient effects of seeded hydrogen on vehicle performance by varying the seed amount during reactor start-up and shut-down events while allowing the engine to use pure hydrogen during steady state operation. This has shown that seeded hydrogen can also benefit the vehicle ΔV performance as a transient had rather than the bulk of the burn profile. It was also suggested that seeded hydrogen could be used to aid in burn events too large for an orbital maneuvering system but too small to justify a full NTP engine burn.
NASA’s Mars Design Reference Architecture 5.0 presents architecture options for a crewed Mars mission. This paper compares the cryogenic chemical-propulsion option in that study with an alternative: water-electrolysis propulsion. Propellant is stored as liquid water instead of cryogenic hydrogen and oxygen, and then electrolyzed on demand into gaseous hydrogen and oxygen for combustion. This addresses a technology gap in the reference architecture: cryogenic propellant storage life. Water is inert and stable, allowing indefinite storage. The mission refuels from pre-positioned tanks at a lunar libration point and Mars orbit, significantly reducing vehicle mass. Low-thrust transfers between Earth escape, Mars encounter, and vice versa reduce the need to store gas for impulsive maneuvers. Only relatively small impulsive maneuvers are used for orbit injection, escape, and plane changes. The proposed architecture achieves the same mission with at most the same number of launch vehicles and without cryogenic propellant storage. Both architectures use five superheavy lift launch vehicles; however, the reference architecture demands all five within 120 days. The alternative is more flexible. Pre-positioning propellant ahead of time allows a less demanding launch cadence. Potential utilization of water from the moon or elsewhere reduces the number of launches from five to as few as two.
As part of a NASA funded effort, Aerojet Rocketdyne, along with other industry partners, has been examining the feasibility of a Low Enriched Uranium (LEU) Nuclear Thermal Propulsion (NTP) system and its application to various human-class Mars missions. Recent analysis for NTP vehicle design focused on Mars opposition class missions in the mid- to late-2030s. After detailed analysis and trade studies, a vehicle configuration was selected that will close on several opposition class missions. These opposition class trajectories require mission Vs in the range of 8-12 km/s, with the lowest case around 6 km/s. These higher Vs require vehicle configurations that are both complex and massive, and involve the development of several NTP elements. An effort began to look at taking the NTP elements designed for opposition class missions, and create a vehicle that could perform conjunction class missions. Mars conjunction class missions typically require 2-3 times less V than opposition class missions do, resulting in lower mass and less complex vehicle configurations. As part of this study, two vehicle configurations were developed with the capability to perform conjunction class missions in the mid- to late- 2030s. Each of these configurations also have excess performance over the minimum energy trajectories, allowing for shorter transit times between Earth and Mars, or additional mission margin. Using the robust set of NTP vehicle elements developed for opposition class missions, two vehicle configurations were designed for conjunction class missions. This paper presents the details and performance of the opposition and conjunction NTP vehicles, along with an introduction to a Mars mission campaign using NTP. Up to this point, analysis on NTP for Mars missions has only evaluated the first crewed mission. Future analysis will focus on the entire set of Mars exploration missions, including the surface logistics and infrastructure required to support a full campaign. While there are several current and advanced propulsion concepts that enable human-class Mars missions, NTP shows that it can be an optimal and robust solution for all Mars missions.
Using the NASA Design Reference Architecture, Nuclear Thermal Propulsion engine and Mars transfer vehicle models were developed to numerically examine the effects of adding heavy noble gases into the hydrogen propellant stream (seeding) on round trip transit times (baseline of 357 days). Seeded hydrogen, up to maximum seed mass concentration (MSMC) 55.85%, increased engine and vehicle performance by reducing pressure losses, decreasing reactor power, and increasing the overall change in velocity while assuming constant vehicle volume and dry mass. The tradeoff was lowered specific impulse and increased net propellant mass, resulting in increased vehicle wetted mass. Vehicle performance increased at MSMC and provided a best case 32-day reduction in transit time vs. pure hydrogen. Vehicle performance was comparable to densified pure hydrogen at 30% seed mass concentration. When taken in combination with densified hydrogen, vehicle performance increased further by providing a 41-day reduction in transit time at MSMC.
Using the NASA Design Reference Architecture, an Aerojet Rocketdyne vehicle architecture, and a University of Alabama in Huntsville nuclear thermal propulsion engine model, the performance of a Mars transfer vehicle using hydrogen propellant seeded with argon was analyzed. Seeded hydrogen, up to a maximum seed mass concentration of 55.85%, increased engine and vehicle performance by reducing pressure losses, decreasing reactor power, and increasing the overall ΔV while assuming constant propellant volume and vehicle dry mass. The tradeoff for using seeded hydrogen was lowered specific impulse and increased net propellant mass, resulting in increased vehicle wetted mass. Vehicle performance monotonically increased with seed mass concentration and provided a best-case 32-day reduction in round-trip transit time to Mars versus pure hydrogen.
If humans are to reach beyond this planet and establish permanent outposts at Mars or any other solar system body, advanced propulsion will be needed. Optimum advanced propulsion needs high thrust to operate within the deep gravity well of a planet and also needs to provide high propulsive efficiency for rapid travel and reduced total spacecraft mass.
One prominent propulsion technology that meets the “optimum” criteria has been researched and tested in the past: it is Nuclear Thermal Propulsion (NTP). NTP, whether designed to provide the thrust to move a spacecraft between orbits or operate as a dual-mode system that provides power and propulsion capability, provides a strong architectural benefit to exploration missions and reusable in-space transportation systems. NTP provides smaller vehicle systems due to its specific impulse (ISP) capability being twice that of the best cryogenic liquid rocket propulsion. The higher ISP also permits reduced trip times when optimizing for the higher delta-V capability for round-trip missions from Earth to Mars. Aerojet Rocketdyne (AR) is working with NASA and other industry partners to improve the design and reduce the cost of NTP engine systems.
AR has been working with NASA and members of industry to perform mission analysis and to design engines and vehicles for Low Enriched Uranium (LEU) approaches for NTP beginning in 2016. LEU can provide a path to a more affordable “new generation” of NTP. 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.
This paper will present the evolution of the conceptual vehicle and engine design performed under the NASA MSFC LEU NTP program. The program has gone through six Design Analysis Cycles (DACs). The mission to be performed is shown and how the mission requirements are translated into vehicle design choices is described. Inputs from various NASA Mars campaign studies and further component studies from NASA were used in the vehicle design. The significant mission analysis that led to the DAC 1 vehicle and engine design is described along with the implications of the analysis results on the individual design choices. The design changes and why they occurred is presented for each of the subsequent DACs.
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 International Space Station is the first space human outpost and over the last 15 years, it has represented a peculiar environment where science, technology and human innovation converge together in a unique microgravity and space research laboratory. With the International Space Station entering the second part of its life and its operations running steadily at nominal pace, the global space community is starting planning how the human exploration could move further, beyond Low-Earth-Orbit. According to the Global Exploration Roadmap, the Moon represents the next feasible path-way for advances in human exploration towards the nal goal, Mars. Based on the experience of the ISS, one of the most widespread ideas is to develop a Cislunar Station in preparation of long duration missions in a deep space environment. Cislunar space is de ned as the area of deep space under the influence of Earth-Moon system, including a set of special orbits, e.g. Earth-Moon Libration points and Lunar Retrograde Orbit. This habitat represents a suitable environment for demonstrating and testing technologies and capabilities in deep space. In order to achieve this goal, there are several crucial systems and technologies, in particular related to transportation and launch systems. The Orion Multi-Purpose Crew Vehicle is a reusable transportation capsule designed to provide crew transportation in deep space missions, whereas NASA is developing the Space Launch System, the most powerful rocket ever built, which could provide the necessary heavy-lift launch capability to support the same kind of missions. These innovations would allow quite-fast transfers from Earth to the Cislunar Station and vice versa, both for manned and unmanned missions. However, taking into account the whole Concept of Operations for both the growth and sustainability of the Cislunar Space Station, the Lunar Space Tug can be considered as an additional, new and fundamental element for the mission architecture. The Lunar Space Tug represents an alternative to the SLS scenario, especially for what concerns all unmanned or logistic missions (e.g. cargo transfer, on orbit assembly, samples return), from Low Earth Orbit to Cislunar space. The paper focuses on the mission analysis and conceptual design of the Lunar Space Tug to support the growth and sustainment of the Cislunar Station. Particular attention is dedicated to the analysis of the propulsion subsystem effects of the Lunar Space Tug design. Main results are presented and discussed, and main conclusions are drawn.
The small modular reactor (SMR) offers many feasible pathways for the construction of more nuclear power plants. A physics model of a near term deployable SMR of the integral pressurized water reactor (IPWR) design is developed. Fuel depletion simulations are performed to optimize the active fuel length, fuel enrichment and core loading pattern in order to achieve a uniform core power distribution. The optimized core can produce 500 MW of thermal power with a four year core life-time at a capacity factor of 87%. The core consists of 69 uranium dioxide (UO2) fuel assemblies; 5 assemblies at 4.4 at% 235U enrichment and 64 assemblies at 4.95 at% 235U enrichment. The active fuel length is 200 cm and the core diameter is 194.55 cm for an active core height-to-diameter ratio of 1.03. As part of the study the active fuel length is increased to 240 cm resulting in an increased capacity factor of 95% at 530 MW of thermal power output for an active core height-to-diameter ratio of 1.23. Rod cluster control assemblies (RCCAs) are placed strategically to reduce the overall core power peaking factor to 1.3. Estimated reactor kinetics parameters such as the delayed neutron fraction and mean neutron generation time are typical of existing larger pressurized water reactors (PWRs) from which much of the IPWR based SMR design is derived. This study showed that Doppler, moderator temperature, void and power reactivity coefficients are all negative over the core life-time of four years indicating the possibility of safe reactor operation. A semi-analytical thermal hydraulics analysis reveals acceptable radial and axial fuel element temperature profiles with significant safety margin from industry standards on peak fuel and clad surface temperature limits. The critical heat flux (CHF) is calculated and is not exceeded even in 10% overpower conditions. In addition the nucleate boiling ratio (DNBR) is calculated and found to be above 4.8 for the entirety of the active core region. These parameters further engender confidence in the safety of the SMR design.
Introduction Experimental Procedure Results and Discussion Summary Acknowledgements
Experimental investigations are conducted on sintered tubular SiC in oxidizing environments containing pure steam at 1 atm with temperature range of 1140-1500 °C and velocity between 0.8 and 10 m/s. Linear weight loss was observed with time. The linear weight loss rates exhibit sensitive dependence on flow rate at a given temperature, demonstrating effects of flow boundary-layer diffusion rate on silica volatilization kinetics. Silica scale exhibits morphology change with respect to exposure time in an oxidizing environment, progressively demonstrating bubble formation and surface smoothing, extensive formation of cracks and pores, and crack reduction. Yet, strength measurements of pressure-less sintered SiC show no significant change after oxidation in the tested conditions. Hence, the primary life-time limiting factor for structural application of pressure-less sintered SiC in the tested environments is anticipated to be the loss of the material. © 2015 Elsevier Ltd and Techna Group S.r.l. All rights reserved.
The Space Shuttle Main Engine (SSME) was the only reusable large liquid rocket engine ever developed. The specific impulse delivered by the staged combustion cycle, substantially higher than previous rocket engines, minimized volume and weight for the integrated vehicle. The dual pre-burner configuration permitted precise mixture ratio and thrust control while the fully redundant controller and avionics provided a very high degree of system reliability and health diagnosis. Power level throttling was required to minimize structural loads on the vehicle early in flight and acceleration levels on the crew late in ascent. Fatigue capability, strength, ease of assembly and disassembly, inspectability, and materials compatibility were all major considerations in achieving a fully reusable design. During the four-decade program the design evolved substantially using a series of block upgrades. A number of materials and manufacturing challenges were encountered throughout SSME’s history. Fracture control was implemented to assess life limits of critical materials and components. Instrumentation systems were a challenge due to the harsh thermal and dynamic environments within the engine. Extensive inspection procedures were developed to assess the engine components between flights. The Space Shuttle Main Engine achieved a remarkable flight performance record. All flights were successful with only one mission requiring an ascent abort condition, which still resulted in an acceptable orbit and mission. This was achieved in large part via extensive ground testing to fully characterize performance and to establish acceptable life limits. During the program over a million seconds of test and flight time were accumulated. Post flight inspections and data assessment were integral to understanding in-flight performance of the hardware. By the end of the program, the predicted reliability had improved by a factor of four. These unique challenges, evolution of the design, and the resulting reliability are discussed in this paper.
A first-order derivation and estimates of the power-permass and specific impulse yield requirements for a nuclear-heated steam rocket (NSR) to deliver payloads to lunar escape from the lunar south pole. The NSR would use lunar-derived water directly as its rocket propellant. Estimates of the ice and cryofuel processing infrastructure masses for the cryofuel alternative provide a reference for comparison. Low Performance, Baseline and High Performance NSR examples calculate payloads delivered and propellant used, given a range of specific power assumptions. Issues of nuclear fuel element temperature and power density are discussed. Results suggest that power density of order 75 megawatts per ton of rocket at a specific impulse of 195 seconds (mixed mean outlet temperature of order 1100 K) are both achievable and will provide highly competitive payloads. For example, a 2 ton, Low Performance NSR tanker making 3 trips per day would deliver of order 240 tons of payload per month to lunar escape. Compared to a cryofueled system, Baseline NSR requires between 150 and 720 times less infrastructure to deliver a given payload to lunar escape. This could imply a dramatic drop in mission cost.
Silicon carbide is a candidate cladding for fission power reactors that can potentially provide better accident tolerance than zirconium alloys. SiC has also been discussed as a host matrix for nuclear fuel. Chemical vapor–deposited silicon carbide specimens were exposed in 0.34–2.07 MPa steam at low gas velocity (~50 cm/min) and temperatures from 1000°C to 1300°C for 2–48 h. As previously observed at lower steam pressure of 0.15 MPa, a two‐layer SiO2 scale was formed during exposure to these conditions, composed of a porous cristobalite layer above a thin, dense amorphous SiO2 surface layer. Growth of both layers depends on temperature, time, and steam pressure. A quantitative kinetics model is presented to describe the SiO2 scale growth, whereby the amorphous layer is formed through a diffusion process and linearly consumed by an amorphous to crystalline phase transition process. Paralinear kinetics of SiC recession were observed after exposure in 0.34 MPa steam at 1200°C within 48 h. High‐pressure steam environments are seen to form very thick (10–100 μm) cristobalite SiO2 layers on CVD SiC even after relatively short‐term exposures (several hours). The crystalline SiO2 layer and SiC recession rate significantly depend on steam pressure. Another model is presented to describe the SiC recession rate in terms of steam pressure when a linear phase transition k l governing the recession kinetics, whereby the reciprocal of recession rate is found to follow a negative unity steam pressure power law.
A brief sketch shows the origin of why and how thermal rocket propulsion has the unique potential to dramatically reduce the cost of space transportation for most inner solar system missions of interest. Orders of magnitude reduction in cost are apparently possible when compared to all processes requiring electrolysis for the production of rocket fuels or propellants and to all electric propulsion systems. An order of magnitude advantage can be attributed to rocket propellant tank factors associated with storing water propellant, compared to cryogenic liquids. An order of magnitude can also be attributed to the simplicity of the extraction and processing of ice on the lunar surface, into an easily stored, non-cryogenic rocket propellant (water). A nuclear heated thermal rocket can deliver thousands of times its mass to Low Earth Orbit from the Lunar surface, providing the equivalent to orders of magnitude drop in launch cost for mass in Earth orbit. Mass includes water ice. These cost reductions depend (exponentially) on the mission delta-v requirements being less than about 6 km/s, or about 3 times the specific velocity of steam rockets (2 km/s, from Isp 200 sec). Such missions include: from the lunar surface to Low Lunar Orbit, (LLO), from LLO to lunar escape, from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO), from LEO to Earth Escape, from LEO to Mars Transfer Orbit, from LLO to GEO, missions returning payloads from about 10{percent} of the periodic comets using propulsive capture to orbits around Earth itself, and fast, 100 day missions from Lunar Escape to Mars. All the assertions depend entirely and completely on the existence of abundant, nearly pure ice at the permanently dark North and South Poles of the Moon. {copyright} {ital 1999 American Institute of Physics.}
Nuclear rocket research and development was initiated in the United States in 1955 and is still being pursued to a limited extent. The major technology emphasis occurred in the decade of the 1960s and was primarily associated with the Rover/NERVA Program where the technology for a nuclear rocket engine system for space application was developed and demonstrated. The NERVA (Nuclear Engine for Rocket Vehicle Application) technology developed twenty years ago provides a comprehensive and viable propulsion technology base that can be applied and will prove to be valuable for application to the NASA Space Exploration Initiative (SEI). This paper, which is historical in scope, provides an overview of the conduct of the NERVA Engine Program, its organization and management, development philosophy, the engine configuration, and significant accomplishments.
This concept proposes to use thermal processes alone to extract water from the lunar South Pole and launch payloads to low lunar orbit. Thermal steam rockets would use water propellant for space transportation. The estimated mass of a space water tanker powered by a nuclear heated steam rocket suggests it can be designed for launch in the Space Shuttle bay. The performance depends on the feasibility of a nuclear reactor rocket engine producing steam at 1,100 degrees Kelvin, with a power density of 150 Megawatts per ton of rocket, and operating for thousands of 20 minute cycles. An example uses reject heat from a small nuclear electric power supply to melt 17,800 tons per year of lunar ice. A nuclear heated steam rocket would use the propellant water to launch and deliver 3,800 tons of water per year to a 100 km low lunar orbit.
Literature thermodynamic values were experimentally confirmed for the
Bunsen reaction producing H2SO4- and HI-rich phases. The sulfur-iodine
water-splitting cycle, which uses the Bunsen reaction, has been improved
by enriching the H2SO4 solution to 57% in a system involving H2SO4
product and counter-current liquid I2 flow. The system was saturated
with SO2. The decomposition of H2SO4 was investigated. Pt/SiO2, Pt/ZrO2,
Pt/TiO2, and Pt/BaSO4 were all shown to be good catalysts for H2SO4
vapor decomposition to SO2 at high temperatures. Pt/Al2O3 was found to
fail due to substrate sulfation. The importance of pressure to sulfation
temperature is presented. A summary of catalyst studies for H2SO4 vapor
decomposition compares catalyst effectiveness.
A new mathematical method is developed for interpolation from a given set of data points in a plane and for fitting a smooth curve to the points. This method is devised in such a way that the resultant curve will pass through the given points and will appear smooth and natural. It is based on a piecewise function composed of a set of polynomials, each of degree three, at most, and applicable to successive intervals of the given points. In this method, the slope of the curve is determined at each given point locally, and each polynomial representing a portion of the curve between a pair of given points is determined by the coordinates of and the slopes at the points. Comparison indicates that the curve obtained by this new method is closer to a manually drawn curve than those drawn by other mathematical methods.
A method is designed for interpolating values given at points of a rectangular grid in a plane by a smooth bivariate function z = z(x, y). The interpolating function is a bicubic polynomial in each cell of the rectangular grid. Emphasis is on avoiding excessive undulation between given grid points. The proposed method is an extension of the method of univariate interpolation developed earlier by the author and is likewise based on local procedures.
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