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In support of NASA’s Unmanned Aircraft Systems Integration in the National Airspace System project and RTCA Special Committee 228, an analysis has been performed to provide insight in to the trade space between detect and avoid (DAA) Well Clear definition threshold variations, which could affect DAA sensor range and alerting requirements.
We measured the impacts on human visual function of a range of vibration levels (0.15, 0.3, 0.5, and 0.7 g) at the frequency and along the axis of the anticipated Ares thrust oscillation. We found statistically significant and equivalent decrements in performance on a reading and a numeric processing task at tested vibration levels above 0.3 g (0-to-peak), but no evidence of after-effects. At the smallest font and highest vibration level tested, the average effect was a 50 percent increase in response time and six-fold increase in errors. Our findings support a preliminary trade space in which currently planned Orion font sizes and text spacing appear to be too small to support accurate and efficient reading at the tested vibration levels above 0.3 g, but not too small to support reading at 0.3 g. This study does not address potential impacts on crew cognitive decision-making or motor control and does not test either the full induced Orion-Ares environment with its sustained Gx-loading or the full complexity of the final Orion seat-helmet-suit interface. A final determination of the Orion-Ares program limit on vibration must take these additional factors into consideration and, thus, may need to be lower than that needed to support effective reading at 1-Gx bias.
This report presents the results of the 2012-2013 NASA Institute for Advanced Concepts (NIAC) Phase 1 "Atom Interferometry for Detection of Gravitational Waves" project. The origin of this GW (Gravitational Wave) detection concept using atoms can be traced to theoretical work that first appeared in 2008 and also to a satellite mission-focused followup study that was done in 2011. The goal of the current project was to explore both theoretical and technical issues surrounding the implementation of this idea, as well as to begin performing proof-of-concept experiments to validate critical aspects of the proposal.The top level trade space for the detector design is driven by the strategy employed to mitigate laser frequency noise, which, if uncontrolled, can mask GW signatures. One of the advantages of the atom interferometric approach is the possibility of single baseline detection (Fig. 1.1), even in the presence of laser noise. This is enabled by the differential measurement between the two ensembles of atoms, which can result in substantial laser noise suppression. The details of this suppression depend on the atomic physics techniques used to implement the atom interferometry. Specifically, we considered the effect on noise suppression that results from using traditional two-photon Raman transitions (with alkali atoms) and also single-photon transitions (with alkaline earth-like atoms).The interferometers shown in Fig 1.1(b) take advantage of single-photon transitions (as opposed to traditional Raman transitions) because using light pulses from one direction at a time allows for near perfect common-mode cancellation of laser phase noise, even for long baselines. This calls for the use of atomic transitions with an (ideally large) optical energy level difference with a long (greater than 1 second) lifetime, such as high-transitions routinely used for optical atomic clocks in species like Sr, Ca and Yb. Notably, large momentum transfer (LMT) atom optics - and the sensitivity enhancement they confer – can still be realized by simply adding additional pairs of alternating pulses to each beam splitter process. Section 3 reports on the theoretical work we performed to justify this GW detection protocol using single-photon transitions. This approach represents a new method for GW detection using atoms that is distinct from the original proposal from 2008. At the system level, we evaluated three architectures, each of which implements a different solution to the laser frequency noise issue. The first two designs are based on two-photon Raman transitions with Rb atoms. One of these is a three-satellite, multiple baseline design while the other is a two-satellite, single baseline design. The third proposal is a two-satellite, single baseline design that uses single-photon transitions with Sr atoms. These three architectures are described in more detail in Section 2. There are a number of known technical issues that we have started to address using ground-based experiments. These issues include atom technology development needs such as, for example, lower ensemble temperature requirements and large momentum transfer (LMT) atom optics. To this end, we have built a 10-meter scale atom drop tower, where we can perform proof-of-principle demonstrations of the proposed AGIS detector in an environment that permits more than 2.5 seconds of free-fall time. This facility allows for demonstration of atom interferometry with long interrogation time (seconds) and large atom wavepacket
In 2015 NASA chartered a partnership between the Jet Propulsion Laboratory (JPL) and the Johns Hopkins Applied Physics Laboratory (APL) to begin planning a mission to study the Jovian moon Europa. The project has adopted a Model-Based Systems Engineering (MBSE) approach to its architecting process since its early formulation, developing certain modeling practices and tools as needed, with the expectation that this process would result in a more consistent and verifiable architecture than with a more traditional document-based approach. A sound architecture is essential to provide the rationale for requirements on the system design, and to define the trade space of acceptable design points within which technical and programmatic concerns as well as project objectives can be addressed. This paper provides an overview of the framework used by the Europa project to describe the mission architecture and discusses how a system model was instrumental in providing a single-source-of-truth for this description. Several key modeling patterns to represent the architecture are presented, along with audit methods to ensure the consistency and the correctness of the model. Finally, the benefits and challenges of using a model-based approach to generate traditional requirements documents and other gate products are assessed.
A potential Mars Sample Return (MSR) mission would require robotic autonomous capture and manipulation of an Orbital Sample (OS) toward returning the samples to Earth. An orbiter would capture the OS, manipulate to a preferential orientation for the samples, transition it through the steps required to break-the-chain with Mars, stowing it in a containment vessel or an Earth Entry Vehicle (EEV) and providing redundant containment to the OS (for example by closing and sealing the lid of the EEV). In this paper, we discuss the trade-space of concepts generated for both the individual aspects of capture and manipulation of the OS as well as concepts for the end-to-end system. Notably, we discuss concepts for OS capture, manipulation of the OS to orient it to a preferred configuration, and steps for transitioning the OS between different stages of manipulation, ultimately securing it in a containment vessel or Earth Entry Vehicle.
A potential Mars Sample Return (MSR) mission could require robotic autonomous capture and manipulation of an Orbital Sample (OS) before returning the samples to Earth. In one scenario, an orbiter would capture the OS, manipulate to a preferential orientation, transition it through the steps required to break-the-chain with Mars, stowing it in a containment vessel or an Earth Entry Vehicle (EEV) and providing redundant containment to the OS (for example by closing and sealing the lid of the EEV). In this paper, we discuss the trade-space of concepts generated for both the individual aspects of capture and manipulation of the OS, as well as concepts for the end-to-end system. Notably, we discuss concepts for OS capture, manipulation of the OS to orient it to a preferred configuration, and steps for transitioning the OS between different stages of manipulation, ultimately securing it in a containment vessel or Earth Entry Vehicle.
Cost is one of the biggest obstacles to sending humans to Mars. However, spacecraft costs are typically not estimated until after the preliminary vehicle and mission concepts have been designed. By automating the cost estimation process, the effect of any change in vehicle or mission design on the mission cost can be determined more efficiently. This paper describes an extension to the tool SAPE (Systems Analysis for Planetary Entry, Descent, and Landing). This extension integrates the cost modeling software SEER-H (System Estimation and Evaluation of Resources-Hardware) with a number of systems analysis tools. This new method is used to analyze several trade spaces of a HIAD (Hypersonic Inflatable Aerodynamic Decelerator) entry vehicle for human Mars missions. Key findings include quantifying the impacts of ballistic coefficient, main engine specific impulse, and thrust-to-weight ratio on the overall cost of the vehicle, and how the payload per lander and number of landers affects the cost of a campaign to Mars.
Enabling safe, routine, and high-density flight operations of small UAS at low-altitude over heavily populated urban centers presents a difficult challenge for emerging UAS Traffic Management (UTM) system concepts. Urban operations by definition involve flight over people, property, and infrastructure. Low-altitude urban environments - such as urban canyons – are one of the most difficult areas for UTM to consider. Mission concepts require routine operations in a cluttered radio-frequency (RF) environment with degraded or denied Global Positioning System (GPS) reception. Flights with any appreciable distance will be beyond visual and communications line-of-sight from ground operators. Timely detection and response to emergencies and onboard failures, which is critical for safe aircraft operation, will be difficult. This work seeks to establish a feasible reference autonomy architecture for autonomous vehicles in an urban UTM system, then verifying and validating this architecture within a complete UTM concept point-design and systems analysis study. In this paper, we present the results from the NASA SAFE50 conceptual design and systems study that investigates the trade-space of urban UTM operations. This advanced conceptual design study develops a feasible, verified, validated point-design solution. The SAFE50 point-design concept places emphasis on advanced, highly-autonomous, and highly-capable vehicles that favors intelligent onboard autonomy over direct human control with today's technologies and operating in today's urban environments. This paper focuses on an general overview of the design study, highlighting decisions made in the architectural solution. This paper will presents a summary of the study, architectures, and requirements. We present an overview of the architecture designs as derived from the top-level UTM system. The point-design has been implemented in both simulation and through flight testing of hardware design prototypes. The results from simulation and flight testing as part of the verification and validation process of the reference design study.
Landsat 9* (L9) is the successor mission to Landsat 8 (L8) previously known as Landsat Data Continuity Mission (LDCM). Both missions are large unmanned remote sensing satellites operating in sunsynchronous polar orbits. As opposed to L8/LDCM, systems engineers for L9 incorporated Micrometeoroid/Orbital Debris (MMOD) protection for small object collisions as part of the L9’s mission success criteria. In other words, the NASA Process for Limiting Orbital Debris (NASA-STD-8719.14A) only calls for analyses of the protection of disposal-critical hardware, but L9 opted to also assess and provide small particle penetration protections for all observatory components including instruments that are not part of the spacecraft components needed for controlled reentry. Systems engineers at Goddard developed a design process to protect against MMOD during the life of Low Earth Orbit (LEO) observatories, and in particular the Landsat 9 Mission. Simply stated, this design process enhanced the effectiveness of existing Multi-Layer Insulation (MLI) to provide the needed protection. The end goal of the design process was to establish a necessary blanket areal density for a given electronics box or instrument wall thickness and a separation between the outer MLI blanket and the structure underneath. The trade space was presented as a set of design curves for different combinations of blanket density, box wall thickness, and separation distance between MLI and structure. An advantage of this process was that it is largely independent of MMOD flux data on a surface-by surface basis. Ultimately, cost savings should result from incorporating small object penetration protection early in the design cycle, rather than adding spot shielding blankets later as needed to meet an overall penetration risk standard (the more traditional approach). The approach and implementation to the L9 Observatory design will be addressed in this paper. *L9 is a joint mission being formulated, implemented, and operated by the National Aeronautics and Space Administration (NASA) and the Department of the Interior’s (DOI) United States Geological Survey (USGS).
The NASA Risk Informed Decision Making process is used to assess a trade space of three dimensionally woven thermal protection systems for application to the Mars Sample Return Earth Entry Vehicle. Candidate architectures are assessed based on mission assurance, technical development, cost, and schedule risk. Assessment methodology differed between the architectures, utilizing a four-point quantitative scale for mission assurance and technical development and highly tailored PERT techniques for cost and schedule. Risk results are presented, in addition to a review of RIDM effectiveness for this application.
The NASA Risk Informed Decision Making process is used to assess a trade space of three dimensionally woven thermal protection systems for application to the Mars Sample Return Earth Entry Vehicle. Candidate architectures are assessed based on mission assurance, technical development, cost, and schedule risk. Assessment methodology differed between the architectures, utilizing a four-point quantitative scale for mission assurance and technical development and highly tailored PERT techniques for cost and schedule. Risk results are presented, in addition to a review of RIDM effectiveness for this application.
Recent studies of human-scale missions to Mars have included a wide trade space of vehicle configurations and control schemes. Some configurations fly at a low angle of attack with a low L/D,while others fly at a high angle of attack with a mid L/D. Some use bank angle control, while others use direct force control, where the angle of attack and sideslip angle are independently modulated. This paper compares the performance of three vehicle configurations: a low-L/D vehicle using direct force control, a low-L/D vehicle using bank control, and a mid-L/D vehicle using bank control. The reference mission is aerocapture at Mars into a highly elliptical, 1-sol orbit. The trajectories are integrated in three degrees of freedom. All three cases utilize numeric predictor-corrector guidances and emulate control system responses with rate and acceleration limits. The configurations are compared using a Monte Carlo analysis. The robustness of each configuration to increased dispersions is also compared.
Cost is one of the biggest obstacles to sending humans to Mars. However, spacecraft costs are typically not estimated until after the preliminary vehicle and mission concepts have been designed. By automating the cost estimation process, the effect of any change in vehicle or mission design on the mission cost can be determined more efficiently. This paper describes an extension to the tool Systems Analysis for Planetary Entry, Descent, and Landing which integrates the cost modeling software System Estimation and Evaluation of Resources-Hardware with a number of systems analysis tools. This new method is used to analyze several trade spaces of an entry vehicle for human Mars missions utilizing a Hypersonic Inflatable Aerodynamic Decelerator and provide preliminary results. Key findings include quantifying how ballistic coefficient, main engine specific impulse, and thrust to weight ratio affect the cost of the vehicle and how the payload per lander and number of landers affects the cost of a campaign to Mars.
Subsonic fixed wing aircraft dominant the world's air fleet and move people and goods on 24/7 basis. While air travel is an effective, efficient means of transportation providing an unmatched combination of speed and range, future subsonic aircraft must continue to improve substantially to meet energy efficiency and environmental compatibility goals. The NASA Fundamental Aeronautics Program via the Subsonic Fixed Wing (SFW) Project addresses the comprehensive challenge of enabling revolutionary energy-efficiency improvements in subsonic transport aircraft combined with dramatic reductions in harmful emissions and perceived noise to facilitate sustained growth of the air transportation system. Advanced technologies and the development of unconventional aircraft systems offer the potential to achieve these improvements. Multidisciplinary advances are required in aerodynamic efficiency to reduce drag, structural efficiency to reduce aircraft empty weight, and propulsive and thermal efficiency to reduce thrust-specific energy consumption (TSEC) for overall system benefit. Additionally, advances are required to reduce perceived noise without adversely affecting drag, weight, or TSEC and to reduce harmful emissions without adversely affecting energy efficiency or noise. The primary focus of SFW is on the "N+3" generation; that is, vehicles that are three generations beyond the current state of the art, requiring mature technology solutions in the 2025-30 timeframe. Success in overcoming these technical challenges will result in major changes to engine cycle and airframe configurations, which in turn will broaden the technology trade space for vehicles ranging from large transports to very light jets. This presentation provides an overview of SFW-sponsored research targeted at improved aerodynamic and structural efficiency for airframe and integrated vehicle systems.
A knowledge-based aerodynamic design method, Constrained Direct Iterative Surface Cur-vature (CDISC), has been leveraged for the computational design exploration of a Common Research Model with Reduced-Sweep and Slotted Wings (CRMRS-SW). Cruise-slotted-wing airfoils include a slot near the trailing edge of a main airfoil component to introduce stream air to the low-momentum, upper-surface boundary layer on a flap airfoil component with the goal of allowing for a more aft loading, relative to supercritical airfoils, without shock-induced separation. Ultimately, the cruise slotted wing is a passive technology that allows for an in-crease in the drag divergence Mach number and/or an increase in aerodynamic efficiency at agiven cruise condition. These aerodynamic benefits may be realized at the vehicle level through a trade space that includes the following design options: increasing cruise Mach number, in-creasing wing thickness, decreasing wing sweep, increasing lift coefficient and/or decreasing drag coefficient. The present study aimed to evaluate the slotted-wing concept’s potential as adrag-saving technology in application to a single-aisle transonic transport aircraft at a given cruise condition. The CDISC design method was used to develop a better understanding of the slotted-wing design variables that have a significant impact on the variability in aircraft drag and leverage this newly-developed knowledge to quantify the aerodynamic performance benefits enabled by the cruise slotted wing.
JPL’s Office of Formulation provides continuity of support and access to domain subject matter experts, as Principal Investigators mature their mission concepts from “cocktail napkin” ideas to Preliminary Design Reviews [1]. Using NASA’s Psyche mission as a case study, we describe JPL’s concurrent engineering A-Team and Team X support to the Psyche competed concept study team in the areas of 1) Initial Feasibility, 2) Trade Space Exploration, 3) Spacecraft Point Design and Cost Estimate, 4) Science, Technical, Management, and Cost Review, and 5) Strategy and Communication Development. NASA’s Psyche Discovery-class mission started as a grassroots idea from Principal Investigator L.T. ElkinsTanton. Is there a compelling Discovery mission to visit the interior of a body for the first time, by sending a mission to an iron metal asteroid? In less than five years the Psyche concept was selected as a mission under NASA’s Discovery Program. While Psyche had a dedicated concept development team [2], they utilized JPL’s concurrent engineering teams, methods, analysis tools, and subject matter experts throughout their mission concept formulation lifecycle
The icy moon oceans beckon with ingredients that potentially may harbour extant life. Beginning with the Galileo and Cassini missions, measurements have revealed the presence of global oceans under the icy crust of several moons of Jupiter and Saturn. Among those moons, Europa and Enceladus have their ocean in contact with the rocky core, providing an environment similar to the conditions existing on the terrestrial seafloor where life has developed at hydrothermal vents. A detailed trade space study was conducted to develop a technology architecture defining a system that would access an icy moon’s ocean. This paper outlines the architecture with specific consideration to the power source necessary to drive the system.
A potential Mars Sample Return (MSR) mission could require robotic autonomous capture and manipulation of an Orbital Sample (OS) before returning the samples to Earth. In one scenario, an orbiter would capture the OS, manipulate to a preferential orientation, transition it through the steps required to break-the-chain with Mars, stowing it in a containment vessel or an Earth Entry Vehicle (EEV) and providing redundant containment to the OS (for example by closing and sealing the lid of the EEV). In this paper, we discuss the trade-space of concepts generated for both the individual aspects of capture and manipulation of the OS, as well as concepts for the end-to-end system. Notably, we discuss concepts for OS capture, manipulation of the OS to orient it to a preferred configuration, and steps for transitioning the OS between different stages of manipulation, ultimately securing it in a containment vessel or Earth Entry Vehicle.