Reference Mission Research Articles

Quantifying the risk of medical evacuation in spaceflight

IntroductionTraditionally, NASA mission planners have used a heuristic and qualitative approach to design medical systems based on prior experience; however, this approach may result in implicit bias in design. The risk of needing to return to definitive care (RTDC) or medevac, has been particularly difficult to quantify due to the complexity of exploration spaceflight. The Informing Mission Planning through Analysis of Complex Tradespaces (IMPACT) tool is a probabilistic risk assessment approach designed by NASA to model medical risk in long duration spaceflight. This paper discusses how this tool can inform RTDC risk and improve the health and safety of astronaut crews. MethodsIMPACT was developed by subject matter experts using an evidence-based approach and can be used to quantify the risk of exceeding the onboard medical capabilities. Within the model, RTDC occurs when a condition meets a specific threshold that exceeds the on-board capabilities and requires a higher level of care. A notional lunar surface design reference mission (DRM) was modeled using IMPACT, and the RTDC rate was analyzed. ResultsOf the 119 medical conditions evaluated, thirty-four were preassigned to have zero probability of RTDC based on lack of acuity. Thirty-one conditions used the need for inpatient hospitalization or admission to intensive care as the RTDC surrogate. Twenty-six conditions used the probability of needing surgery, and ten conditions used treatment failure rates as the RTDC surrogate. The remaining 18 conditions had unique surrogates that did not fall into the above categories or were assigned a 100 % probability of RTDC due to their definition. In the sample DRM, the overall risk for RTDC was 0.32 events per mission. DiscussionQuantifying medical risk for human spaceflight is challenging but important in designing medical systems that support the health and performance of our astronaut crews. IMPACT can use the best available evidence to quantify the risk of an RTDC event based on mission parameters and help mission planners make informed design decisions. More work is underway to further refine the model and evidence that drives the RTDC calculation to improve model fidelity and utility for missions to the Moon, Mars, and beyond.

Space-Based Passive Orbital Maneuver Detection Algorithm for High-Altitude Situational Awareness

Orbital maneuver detection for non-cooperative targets in space is a key task in space situational awareness. This study develops a passive maneuver detection algorithm using line-of-sight angles measured by a space-based optical sensor, especially for targets in high-altitude orbit. Emphasis is placed on constructing a new characterization for maneuvers as well as the corresponding detection method. First, the concept of relative angular momentum is introduced to characterize the orbital maneuver of the target quantitatively, and the sensitivity of the proposed characterization is analyzed mathematically. Second, a maneuver detection algorithm based on the new characterization is designed in which sliding windows and correlations are utilized to determine the mutation of the maneuver characterization. Subsequently, a numerical simulation system composed of error models, reference missions and trajectories, and computation models for estimating errors is established. Then, the proposed algorithm is verified through numerical simulations for both long-range and close-range targets. The results indicate that the proposed algorithm is effective. Additionally, the sensitivity of the proposed algorithm to the width of the sliding window, accuracy of the optical sensor, magnitude and number of maneuvers, and different relative orbit types is analyzed, and the sensitivity of the new characterization is verified using simulations.

ROV-Based Autonomous Maneuvering for Ship Hull Inspection with Coverage Monitoring

Hull inspection is an important task to ensure sustainability of ships. To overcome the challenges of hull structure inspection in an underwater environment in an efficient way, an autonomous system for hull inspection has to be developed. In this paper, a new approach to underwater ship hull inspection is proposed. It aims at developing the basis for an end-to-end autonomous solution. The real-time aspect is an important part of this work, as it allows the operators and inspectors to receive feedback about the inspection as it happens. A reference mission plan is generated and adapted online based on the inspection findings. This is done through the processing of a multibeam forward looking sonar to estimate the pose of the hull relative to the drone. An inspection map is incrementally built in a novel way, incorporating uncertainty estimates to better represent the inspection state, quality, and observation confidence. The proposed methods are experimentally tested in real-time on real ships and demonstrate the applicability to quickly understand what has been done during the inspection.

Lunar SmallSat Missions with Chemical-Electrospray Multimode Propulsion

Multimode spacecraft propulsion combines two or more propulsive modes into a single system using a single propellant. In this study, a multimode system combining a monopropellant chemical decomposition mode with an electric electrospray mode is considered for specific NASA-relevant lunar small satellite missions. This multimode concept is compared to all-chemical and all-electric approaches for four design reference missions (DRMs) in terms of feasibility, propellant mass consumed, and transfer time. Point solution trajectories are developed using NASA’s General Mission Analysis Tool for each propulsion system. This is the first study to consider this multimode concept for specific missions with high-fidelity force modeling. The results demonstrate the multimode approach can complete three of the four DRMs within the allotted propellant budget, while the all-chemical approach fails to complete two DRMs and the all-electric system fails to complete one DRM. Across the three feasible DRMs, the multimode approach completes the mission in 11–55% less time than the all-electric approach while consuming up to 33% less propellant than the all-chemical approach. The number of switching events and the minimum required mode switching time are quantified for the first time, with the latter having values between nearly instantaneous and 112.5 days, depending on the mission.

Cost estimation for launch vehicle families considering uncertain market scenarios

When optimizing the staging of a launch vehicle, usually a reference mission with a fixed payload mass is used to determine the specific launch cost (or a surrogate thereof, e.g. total mass or dry mass). While this approach neglects suboptimal loading, it delivers sufficiently good results for the evaluation of an individual launcher if the design mission is well known. However, this approach does not capture the full range of payloads that a launch vehicle is expected to launch during its operational lifetime. In order to estimate the cost and subsequently optimize the staging of such launch vehicle families, a full launch scenario has to be considered. The cost of an individual launch vehicle loses importance and the deciding factor is the total cost associated with launching the specified scenario. However, estimates of the future market inevitably include large uncertainties. These uncertainties and their impact are ideally considered and quantified when assessing the specific recurring cost of a single launch vehicle. The consideration of a whole launch market scenario is especially relevant when evaluating a family of launch vehicles, an option currently being discussed for future European launchers. In this case, the distribution of the payloads onto the different members of the launch vehicle family has to be optimized in order to reach the lowest possible cost. Within this manuscript, a method that allows the determination of the recurring cost of such a launcher family based on a launch market scenario including uncertainties is proposed. Based on the launch numbers derived from the combinatorial optimization, parametric cost models for recurring and non-recurring cost are adapted and applied to three possible future launch vehicle families.

CryoSat Long-Term Ocean Data Analysis and Validation: Final Words on GOP Baseline-C

ESA’s Earth explorer mission CryoSat-2 has an ice-monitoring objective, but it has proven to also be a valuable source of observations for measuring impacts of climate change over oceans. In this paper, we report on our long-term ocean data analysis and validation and give our final words on CryoSat-2’s Geophysical Ocean Products (GOP) Baseline-C. The validation is based on a cross comparison with concurrent altimetry and with in situ tide gauges. The highlights of our findings include GOP Baseline-C showing issues with the ionosphere and pole tide correction. The latter gives rise to an east–west pattern in range bias. Between Synthetic Aperture Radar (SAR) and Low-Resolution Mode (LRM), a 1.4 cm jump in range bias is explained by a 0.5 cm jump in sea state bias, which relates to a significant wave height SAR-LRM jump of 10.5 cm. The remaining 0.9 cm is due to a range bias between ascending and descending passes, exhibiting a clear north–south pattern and ascribed to a timing bias of +0.367 ms, affecting both time-tag and elevation. The overall range bias of GOP Baseline-C is established at −2.9 cm, referenced to all calibrated concurrent altimeter missions. The bias drift does not exceed 0.2 mm/yr, leading to the conclusion that GOP Baseline-C is substantially stable and measures up to the altimeter reference missions. This is confirmed by tide gauge comparison with a selected set of 309 PSMSL tide gauges over 2010–2022: we determined a correlation of R = 0.82, a mean standard deviation of σ=5.7 cm (common reference and GIA corrected), and a drift of 0.17 mm/yr. In conclusion, the quality, continuity, and reference of GOP Baseline-C is exceptionally good and stable over time, and no proof of any deterioration or platform aging has been found. Any improvements for the next CryoSat-2 Baselines could come from sea state bias optimization, ionosphere and pole tide correction improvement, and applying a calibrated value for any timing biases.

Mars mission capabilities enabled by nuclear thermal propulsion

As part of the Artemis era of space exploration, space agencies will be working together with their industry partners to establish systems and infrastructure that enable sustained Lunar missions and develop capabilities for Mars. Lockheed Martin has a longstanding history of designing, building, and operating systems for deep space applications, from Viking to Orion. Lockheed Martin develops concepts, trades, and early system designs for Moon and Mars exploration to enable future missions. In 2016, Lockheed Martin presented a vision for achieving crewed exploration of Mars. Known as Mars Base Camp (MBC), this design reference mission envisioned a crewed vehicle in Martian orbit from which astronauts could perform excursions to Phobos and Deimos and perform telerobotic exploration of the Martian surface, including sample return. This concept served as an “existence proof” for a novel, practical, and affordable path to enable human exploration of the Martian system. In 2017, an update was published that included the production of propellant from water and a single-stage fully reusable Mars lander concept. Then in 2021, Lockheed Martin introduced a nuclear thermal propulsion (NTP) configuration of Mars Base Camp, a technology that will be greatly enabling for crewed missions to Mars. This paper matures the design of the NTP configuration of Mars Base Camp and explores the Mars mission capabilities that the NTP technology will enable. These investigations include optimized trajectories for faster transit to Mars and an analysis of abort scenarios that are made possible by the high thrust and performance of NTP. The results provide insight into the performance and capabilities that are realized when NTP technology is matured and integrated into a Mars transit vehicle. The paper will also describe initial surface infrastructure capabilities that significantly enhance sortie science operations, including mobility and power. The resulting updated Mars Base Camp serves as an example of what is possible with current technologies and technologies under development. The MBC reference mission builds on the capabilities that will be developed and demonstrated through Artemis missions, enabling a viable, sustainable long-term Mars exploration program, with the first crewed Mars mission possible in the 2030s.

Lunar Orbit acquisition of the Korea Pathfinder lunar orbiter: Design reference vs actual flight results

The Korea Pathfinder Lunar Orbiter (KPLO), officially named as Danuri, is the Republic of Korea's first spacecraft to orbit the Moon. The goal of KPLO is to secure core deep space technology for future space exploration in Korea and contribute to lunar science by gathering data from the lunar orbit for one year. This paper presents the results of KPLO's Lunar Orbit Acquisition (LOA) phase in detail, with a particular focus on the operational results of Flight Dynamics (FD). During the LOA phase, the KPLO FD team experienced many expected and unexpected issues. These issues were mitigated by making real-time updates to the LOA phase Design Reference Mission (DRM). The paper also presents the results of the Orbit Determination (OD) and performance of each Lunar Orbit Insertion (LOI) maneuver conducted during the actual flight operation. The importance of considering operational constraints when designing the DRM is emphasized by presenting lessons learned based on actual flight experiences. Unlike the original KPLO LOA phase DRM, KPLO successfully achieved the final mission orbit using only three LOI burns, instead of the planned five LOI burns and an Orbit Trim Maneuver (OTM). Despite the extensive revisions made to the DRM, all mission requirements were still fulfilled during the final lunar orbit insertion, taking into account the significant operational constraints.

Selecting Medical Conditions Relevant to Exploration Spaceflight to Create the IMPACT 1.0 Medical Condition List.

INTRODUCTION: Medical conditions occurring in spaceflight pose risks to the crew and the mission and these risks will be exacerbated during exploration-class missions. Probabilistic risk assessment is a method used at NASA to quantify this risk for low-Earth orbit operations. Informing Mission Planning via Analysis of Complex Tradespaces (IMPACT) is a next-generation tool suite that will perform these assessments for exploration-class missions. It will require a robust list of medical conditions of significant likelihood and/or consequence to exploration-class missions to accurately inform the tool suite.METHODS: The IMPACT 1.0 Medical Condition List (ICL 1.0) contains 120 conditions selected in the context of a 210-d cis-lunar, Mars analog design reference mission. The conditions were selected via a systematic process that preserved institutional knowledge from nine prior condition lists. Conditions were prioritized for inclusion in the ICL 1.0 based on history of occurrence in spaceflight, concurrence among the nine source lists, and concurrence among subject matter experts.DISCUSSION: The ICL 1.0 has notable advantages over its predecessor lists in that it is more specific to exploration-class missions, contains a greater number, breadth, and depth of conditions, and was derived via consensus across multiple medical specialties.Kreykes AJ, Suresh R, Levin D, Hilmers DC. Selecting medical conditions relevant to exploration spaceflight to create the IMPACT 1.0 Medical Condition List. Aerosp Med Hum Perform. 2023; 94(7):550-557.

Risk trade-space analysis for safe human expeditions to Mars

We assessed the integrated safety, health, and performance risk to crews on long-duration missions, specifically to Mars. Using a systems approach rather than one focused on individual countermeasures, we examined the trade space around several such risks to identify high-potential risk mitigation strategies and characterize aspects of Mars mission architectures that could lower aggregated risk. Current Mars Design Reference missions would require durations well over two years and would increase crew exposure to radiation and microgravity well beyond ISS levels, likely resulting in significantly reduced performance beyond our current capability to mitigate that could jeopardize mission success. A “fast Mars transit” round-trip mission concept was studied using an innovative flight dynamics approach to quantify the minimum total mission energy required for a Mars transit with total mission duration less than 400 days. This approach holds promise for sending humans to Mars and returning them safely with acceptable, potentially mitigatable, exposure to microgravity and radiation using current or near-term technologies. The fast transit concept would also result in fewer time-driven vehicle failures and enable sustainable deployment of humans and infrastructure to Mars on a regular cadence, allowing steady exploration and colonization of Mars. Finally, we conclude that reliance on the Low Earth Orbit (LEO) mission operations paradigm – i.e., one of near-complete real-time dependence on experts at Mission Control to manage the combined state of the mission, vehicle, and crew – is high risk given the communication delays and limited resupply of any Mars mission, and this risk is not eliminated by the shorter missions durations of fast transit scenarios. Based on historical trends, it is highly likely that the crew will face a high-consequence problem of uncertain origin during Mars transit when ground support will be greatly reduced. While it may be possible to reduce anomaly rates through improved reliability analysis and testing, and to reduce anomaly impacts through added robustness, such mitigations address only known failure modes and known uncertainties. Therefore, a radical shift in the Human-Systems Integration Architecture (HSIA) that defines the operational paradigm, systems design, and human-systems interactions is required to improve the risk posture to an acceptable level regardless of mission duration.

Current observed global mean sea level rise and acceleration estimated from satellite altimetry and the associated measurement uncertainty

Abstract. We present the latest release of the global mean sea level (GMSL) record produced by the French space agency Centre National d’Etudes Spatiales (CNES) and distributed on the AVISO+ website. This dataset is based on reprocessed along-track data, so-called L2P 21, of the reference missions TOPEX/Poseidon (TP) and Jason-1, Jason-2 and Jason-3. The L2P 21 CNES/AVISO+ GMSL record covers the period January 1993 to December 2021 and is now delivered with an estimate of its measurement uncertainties following the method presented in Ablain et al. (2019). Based on the latest calibration (Cal) and validation (Val) knowledge, we updated the uncertainty budget of the reference altimetry mission measurements and demonstrate that the CNES/AVISO+ GMSL record now achieves stability of performances of ± 0.3 mm yr−1 at the 90 % confidence level (C.L.) for its trend and ±0.05 mm yr−2 (90 % C.L.) for its acceleration over the 29 years of the altimetry record. Thanks to an analysis of the relative contribution of each measurement uncertainty budget contributor, i.e. the altimeter, the radiometer, the orbit determination and the geophysical corrections, we identified the current limiting factors to the GMSL monitoring stability and accuracy. We find that the radiometer wet troposphere correction (WTC) and the high-frequency errors with timescales shorter than 1 year are the major contributors to the GMSL measurement uncertainty over periods of 10 years (30 %–70 %), for both the trend and acceleration estimations. For longer periods of 20 years, the TP data quality is still a limitation, but more interestingly, the International Terrestrial Reference Frame (ITRF) realization uncertainties becomes dominant over all the other sources of uncertainty. Such a finding challenges the altimetry observing system as it is designed today and highlights clear topics of research to be explored in the future to help the altimetry community to improve the GMSL measurement accuracy and stability.

Space vehicle docking system standardization

The International Docking System Standard (IDSS) was developed and establish to aid on-orbit crew rescue and joint operations between different spacecraft. For the International Space Station (ISS), the IDSS has successfully enabled Global interoperability for Commercial Crew and it is now being extended to the Artemis campaign. Similarly, as more companies, agencies, and nations announce their intentions to explore and occupy Low Earth Orbit (LEO) and Cis-Lunar space, including the Lunar surface, it is a natural supposition that new, vehicle interface standards will be required to support the build-up of infrastructure for campaign-based exploration or permanent occupation by national and multi-national Agencies, Industries, and Companies.A surface version of the IDSS, a.k.a. IDSS-Surface (IDSS-S), is under consideration at the NASA Johnson Space Center (JSC) by the docking discipline leads responsible for the leadership of technical development and negotiation of the original IDSS over a decade ago. The IDSS-S, like its predecessor, will ultimately detail the physical geometric mating interface and design load requirements to ensure physical interoperability and to support a broad set of design reference missions which, if accommodated, increases the probability of successful Lunar surface docking between different modules enabling the accessibility and inclusivity required for multi-national, sustainable Lunar exploration.

Analysis of cost-efficient urban air mobility systems: Optimization of operational and configurational fleet decisions

With the introduction of low-noise and low-emission electric vertical take-off and landing vehicles, passenger air transportation in urban areas is becoming increasingly important. Previous studies have designed vehicle concepts based on reference mission profiles, however, without considering strategic decisions about fleet operations, such as charging infrastructure, range and battery capacity. Comprehensive cost analyses for urban air mobility fleets have been largely neglected. In this paper, we develop an optimization model to analyze the cost-efficient operation of urban air mobility systems. Strategic decisions on vehicle concepts, battery capacity, and charging infrastructure are incorporated and evaluated using a total cost of ownership approach. We consider state-of-the-art modeling approaches, including vehicle-specific parameters, for accurate calculation of energy consumption. The optimization model is applied to the multi-agent transport simulation MATSim in a case study of the Ruhr (Germany) scenario to evaluate key parameter variations.The study results indicate the feasibility of an urban air mobility system in the study region with trip costs ranging from 27 US-$ to 46 US-$ per requested trip, depending on the scenario settings. Based on parameter variations, we find that a high cruising speed has a detrimental effect on the total cost. A medium charging power level is sufficient and energy costs account for only a moderate share. The latter is in contrast to previous research results, but can be attributed to the more detailed modeling approach of vehicle-specific parameters. We propose a cost incentive system as the ground handling time of the vehicles outweighs the recharging and flight time. The overall results provide a promising basis for model extensions.

Preliminary Design and Simulation of a Thermal Management System with Integrated Secondary Power Generation Capability for a Mach 8 Aircraft Concept Exploiting Liquid Hydrogen

This paper introduces the concept of a thermal management system (TMS) with integrated on-board power generation capabilities for a Mach 8 hypersonic aircraft powered by liquid hydrogen (LH2). This work, developed within the EU-funded STRATOFLY Project, aims to demonstrate an opportunity for facing the challenges of hypersonic flight for civil applications, mainly dealing with thermal and environmental control, as well as propellant distribution and on-board power generation, adopting a highly integrated plant characterized by a multi-functional architecture. The TMS concept described in this paper makes benefit of the connection between the propellant storage and distribution subsystems of the aircraft to exploit hydrogen vapors and liquid flow as the means to drive a thermodynamic cycle able, on one hand, to ensure engine feed and thermal control of the cabin environment, while providing, on the other hand, the necessary power for other on-board systems and utilities, especially during the operation of high-speed propulsion plants, which cannot host traditional generators. The system layout, inspired by concepts studied within precursor EU-funded projects, is detailed and modified in order to suggest an operable solution that can be installed on-board the reference aircraft, with focus on those interfaces impacting its performance requirements and integration features as part of the overall systems architecture of the plane. Analysis and modeling of the system is performed, and the main results in terms of performance along the reference mission profile are discussed.

Task impairment: A novel approach for assessing impairment during exploration-class spaceflight missions

Long duration deep space exploration missions require an innovative assessment for medical impairment. Task Impairment is a novel, dynamic, and mission-appropriate impairment paradigm which will replace Functional Impairment, the metric used currently by the NASA Integrated Medical Model for probabilistic risk assessment on the International Space Station. To derive Task Impairment, the Human Exploration of Mars: Preliminary List of Crew Tasks was used as the source of likely exploration mission tasks. Tasks were divided into one or more human system task categories (e.g., cardiopulmonary, cognitive, etc.). Subject matter experts across five medical specialties then reviewed medical conditions from the Informing Mission Planning via Analysis of Complex Tradespaces condition list to determine which of the 18 human system task categories were impaired in best and worst-case scenarios for treated and untreated variants of each condition. If a human system task category was determined to be affected by a condition, every task assigned to that category was presumed to be, at least partially, impaired. The resulting total tasks impaired by each medical condition were used to calculate Task Impairment values. As a direct assessment of a crewmember's ability to perform mission specific tasks, Task Impairment has the capacity to yield a higher fidelity measure of reduced crew ability than previous techniques. This technique is easily adapted to future proposed design reference missions and may be useful in defining mission phase-specific disability as well as a future medical loss of mission objectives metric.

Emission-Driven Hybrid Rocket Engine Optimization for Small Launchers

Hybrid rocket engines are a green alternative to solid rocket motors and may represent a low-cost alternative to kerosene fueled rockets, while granting performance and control features similar to that of typical storable liquid rocket engines. In this work, the design of a three-stage hybrid launcher is optimized by means of a coupled procedure: an evolutionary algorithm optimizes the engine design, whereas an indirect optimization method optimizes the corresponding ascent trajectory. The trajectory integration also provides the vertical emission profiles required for the evaluation of the environmental impact of the launch. The propellants are a paraffin-based wax and liquid oxygen. The vehicle is launched from the ground and uses an electric turbo pump feed system. The initial mass is given (5000 kg) and the insertion of the payload into a 600-km circular, and polar orbit is considered as a reference mission. Clusters of similar hybrid rocket engines, with only few differences, are employed in all stages to reduce the development and operational costs of the launcher. Optimization is carried out with the aim of maximizing the payload mass and then minimizing the overall environmental impact of the launch. The results show that satisfactory performance is achievable also considering rocket polluting emissions: the carbon footprint of the launch can be reduced by one fourth at the cost of a 5-kg payload mass reduction.

Deep-neural-network-based angles-only relative orbit determination for space non-cooperative target

Angles-only relative orbit determination for space situational awareness is subjected to a well-known range observability problem, that is, the range between two spacecraft is weakly observable or unobservable. However, in previous studies, the solutions to this problem using nonlinear relative motion dynamics have been limited by the computational complexity and long arc of observation. To this end, this study develops a simple and fast algorithm to address the angles-only relative orbit determination problem by exploiting a deep neural network (DNN), which has a significantly strong capability of capturing nonlinearity. Emphasis is placed on the construction of a nonlinear mapping model from the line-of-sight angles to the relative orbit state by training the designed DNN. First, a training dataset generator including nonlinear relative motion dynamics and a line-of-sight measurement model is established to generate the training data for the DNN. Second, the DNN frame, including the network structure, data processing, and network training algorithm for angles-only relative orbit determination, is designed. Subsequently, a digital simulation system comprising error models, reference missions and trajectories, and computation models for error estimation is established. Thus, the anti-noise and generalization performance of the nonlinear mapping model on GEO-type orbits are verified using digital simulations. The results indicate that the proposed algorithm is effective, whereby the estimation accuracy for the relative position is generally better than that for the relative velocity. In the case of a co-elliptical orbit, the estimated errors of distance and velocity in each direction are less than 9.7% and 48.2%, respectively, whereas the maximum average errors are approximately 1.1% and 4.2%, respectively, where only three sets of angle measurements with an interval of 600 s are available. However, in the case of a non-coplanar orbit, the interval between angles can be decreased to 50 s, when the corresponding estimated error of distance in each direction is less than 9.9%, and the maximum average error is approximately 1.1%. Additionally, the sensitivity of the proposed algorithm to the arc length, number, and interval of the measurements is analyzed.

Heliocentric Effects of the DART Mission on the (65803) Didymos Binary Asteroid System

The Double Asteroid Redirect Test (DART) is NASA’s first kinetic impact–based asteroid deflection mission. The DART spacecraft will act as a projectile during a hypervelocity impact on Dimorphos, the secondary asteroid in the (65803) Didymos binary system, and alter its mutual orbital period. The initial momentum transfer between the DART spacecraft and Dimorphos is enhanced by the ejecta flung off the surface of Dimorphos. This exchange is characterized within the system by the momentum enhancement parameter, β, and on a heliocentric level by its counterpart, β ⊙. The relationship between β and the physical characteristics of Dimorphos is discussed here. A nominal set of Dimorphos physical parameters from the design reference asteroid and impact circumstances from the design reference mission are used to initialize the ejecta particles for dynamical propagation. The results of this propagation are translated into a gradual momentum transfer onto the Didymos system barycenter. A high-quality solar system propagator is then used to produce precise estimates of the post-DART encounters between Didymos and Earth by generating updated close approach maps. Results show that even for an unexpectedly high β ⊙, a collision between the Didymos system and Earth is practically excluded in the foreseeable future. A small but significant difference is found in modeling the overall momentum transfer when individual ejecta particles escape the Didymos system, as opposed to imparting the ejecta momentum as a single impulse at impact. This difference has implications for future asteroid deflection campaigns, especially when it is necessary to steer asteroids away from gravitational keyholes.

Simultaneous aircraft sizing and multi-objective optimization considering off-design mission performance during early design

Traditional aircraft conceptual design primarily involves determination of top-level sizing parameters, resulting in an initial design which satisfies specified point-performance constraints while flying the so-called design mission. In practical scenarios, commercial aircraft are also expected to operate optimally in the actual missions they fly which may drastically differ from the design mission. To improve performance and reduce operating costs, optimization shall be performed using specific objectives from the on-design mission and from one or more representative off-design reference mission(s). Such multi-mission optimizations may result in different designs for the same performance constraints. Moreover, the size of aircraft and choice of reference mission(s) may also have an effect in the difference resulting from such multi-mission optimizations. This paper solves a series of multi-objective on-design and multi-mission optimization problems in the conceptual design phase on aircraft spanning a range of size-classes, using the Non-dominated Sorting Genetic Algorithm II (NSGA-II). Results shed light on the design differences between formulations that exclusively consider design mission metrics of interest from the ones that consider metrics of interest from disparate, as-flown, i.e., off-design missions. In addition, results also reveal the impact of off-design mission weightings on the designs obtained from the optimization problems.

Development and Evaluation of Control Architectures for a Mechanically Deployed Entry Vehicle

The NASA-funded Pterodactyl Project seeks to advance the current state of the art for atmospheric entry vehicles by developing novel guidance and control technologies for deployable entry vehicles (DEVs). This paper details the efforts of the Pterodactyl team to develop, implement, and evaluate different control hardware effectors on an asymmetric, mechanically deployed DEV. Multiple control effector designs are developed for a baseline vehicle including propulsive control systems (reaction control systems) and nonpropulsive control systems (aerodynamic control surfaces and internal moving masses). For each control system, state-feedback integral controllers based on linear quadratic regulator optimal control methods are designed to track the guidance commands of either the bank angle or the angle of attack and sideslip angle. Using a lunar return reference mission, a comparative analysis of the performance and maneuverability of the DEV using the different control system configurations is conducted. The aerodynamic control surface design is demonstrated to be the most suitable architecture, primarily due to its ability to achieve acceleration rates that supersede the other designs. Considering performance, maneuverability, and guidance tracking during entry, the aerodynamic control surface design is recommended for this DEV and chosen reference mission.