NASA Engineering and Safety Center

Almost fifty years ago, Dr. Robert Frosch, the fifth NASA administrator, delivered a speech on the practice of systems engineering and a call for improvement that remains relevant today. A key difference between the past and today is that the digital information revolution offers new ways for improvements to manage complex systems, enhance knowledge transfer, and improve communications. These new ways offer reductions in cost, schedule, and risk. Over the past several years, the NASA Systems Engineering community has explored ways to use the digital environment to realize these benefits. The NASA Systems Engineering community began evaluating the adoption of a digital approach or Model-Based Systems Engineering (MBSE) as early as 2011. The effort performed benchmarking of industry, evaluated standards, discussed infrastructure requirements, and interviewed NASA stakeholders. In 2016, the MBSE Pathfinder was established to evaluate application of MBSE to some of the most challenging aspects of real NASA spaceflight systems. The following year, the MBSE Pathfinder expanded to more rigorous implementation and coverage across multiple phases of the system engineering lifecycle. Two years of learning, alignment, and application resulted in over a dozen concrete use cases that illustrate the benefits of a digital framework for systems engineering. The MBSE Pathfinder informed the plans to move NASA towards enterprise implementation of MBSE. This paper highlights examples in which both quantitative and qualitative benefits were obtained. The first example shows how systems engineering models used for concept design and definition were re-used for verification and test for a rocket engine. Schedule metrics were tracked to determine the improvement when using MBSE compared to recent historical data from manual methods. The second example shows the seamless transfer of modeling elements and data between the systems engineering model, computer aided design model, loads analysis, and additive manufacturing software for a payload adapter. This example shows how MBSE can be used to validate concepts and perform rapid prototyping. Qualitative benefits include improved communications and customer satisfaction. These examples, and others in a rich portfolio of results, demonstrate benefits available at points across the entire lifecycle in a rapid, agile approach. The MBSE Pathfinder projects were set up deliberately to sample multiple points in the lifecycle to understand the nuances and requirements sooner, rather than following one project end-to-end. The results from the MBSE Pathfinder help define the infrastructure and requirements necessary for a future systems engineering capability and provide a guide towards a full-up, integrated approach to systems development. To organize and implement these goals, the MBSE Pathfinder expanded into the MBSE Infusion and Modernization Initiative (MIAMI). The approach as well as the structure of MIAMI is unique in that it relies on learning from the use of MBSE on real projects and capturing the benefits that are useful for working engineers to make systems engineering easier. The lessons from this work are available for use by those who are interested in learning from or partnering with MIAMI.

Wind tunnel tests were conducted by Nielsen Engineering & Research (NEAR) and Rose Engineering & Research (REAR) in conjunction with the NASA Engineering & Safety Center (NESC) on a 6%-scale model of the Orion launch abort vehicle (LAV) configured with four grid fins mounted near the base of the vehicle. The objectives of these tests were to 1) quantify LAV stability augmentation provided by the grid fins from subsonic through supersonic Mach numbers, 2) assess the benefits of swept grid fins versus unswept grid fins on the LAV, 3) determine the effects of the LAV abort motors on grid fin aerodynamics, and 4) generate an aerodynamic database for use in the future application of grid fins to small length-to-diameter ratio vehicles similar to the LAV. The tests were conducted in NASA Ames Research Center's 11x11-foot transonic wind tunnel from Mach 0.5 through Mach 1.3 and in their 9x7-foot supersonic wind tunnel from Mach 1.6 through Mach 2.5. Force- and moment-coefficient data were collected for the complete vehicle and for each individual grid fin as a function of angle of attack and sideslip angle. Tests were conducted with both swept and unswept grid fins with the simulated abort motors (cold jets) off and on. The swept grid fins were designed with a 22.5deg aft sweep angle for both the frame and the internal lattice so that the frontal projection of the swept fins was the same as for the unswept fins. Data from these tests indicate that both unswept and swept grid fins provide significant improvements in pitch stability as compared to the baseline vehicle over the Mach number range investigated. The swept fins typically provide improved stability as compared to the unswept fins, but the performance gap diminished as Mach number was increased. The aerodynamic performance of the fins was not observed to degrade when the abort motors were turned on. Results from these tests indicate that grid fins can be a robust solution for stabilizing the Orion LAV over a wide range of operating conditions.

The model for aeroelastic validation research involving computation semispan wind-tunnel model, a transport wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental limit cycle oscillation behavior data at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wing-tip configurations: clean, tip store, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6-0.9, and maps of limit cycle oscillation behavior were made in the range of M = 0.85-0.95. The effects of dynamic pressure and angle of attack were measured. Testing in both R134a heavy gas and air provided unique data on the Reynolds number, transition effects, and the effect of speed of sound on limit cycle oscillation behavior. This report gives an overview of the test results, including experimental flutter boundaries, and the conditions involving shock-induced transonic flow separation onset at low wing angles, including maps of limit cycle oscillation behavior.

Abstract In 2016, the NASA Engineering and Safety Center established a model‐based systems engineering (MBSE) Pathfinder. The primary motivations for establishing the MBSE Pathfinder were to advance the Agency's applications of MBSE and capture lessons‐learned to inform the next steps. The MBSE Pathfinder had four teams working in parallel for eight months on different topics of interest to NASA. The teams were encouraged to learn, and use creativity and innovation in their system modeling. The results were captured via reports, webinars, and a knowledge capture meeting. The approach taken for the MBSE Pathfinder was very successful in providing a number of lessons‐learned for NASA and for other organizations considering MBSE or pathfinder efforts, and in building a very strong and collaborative user community.

The agility of a rigid-body spacecraft can be expressed in terms of a geometric, three-dimensional, solid called the agilitoid. Originally developed as a means for explaining the concept of “hidden agility” made visible through the use of optimal control techniques, a modified agilitoid called an agility envelope is presented here that is compatible with conventional eigenaxis maneuvers. This paper demonstrates how the agility envelope can be applied to size an attitude control system (ACS) and/or assess the capability of an existing design. Analysis of the James Webb Space Telescope (JWST) ACS shows that the agility envelope accurately predicts the true capability of the ACS: a 90 deg maneuver can actually be completed 15% faster than the conventional back-of-the-envelope slew-sizing equations suggest. The utility of the agility envelope is further illustrated by showing how an alternative control allocation scheme can reduce the JWST torque and momentum requirements by 40%. The otherwise hidden agility can be recovered to enhance the slew performance of the JWST or allow the reaction wheel array to be reduced from six to five wheels, while meeting existing maneuver requirements. The agility envelope allows such design trades to be studied without the need to perform detailed simulations of the attitude control system.

The present paper describes the structural analyses performed on a preloaded bolted-joint configuration. The joint modeled was comprised of two L-shaped structures connected together using a single bolt. Each L-shaped structure involved a vertical flat segment (or shell wall) welded to a horizontal segment (or flange). Parametric studies were performed using elasto-plastic, large-deformation nonlinear finite element analyses to determine the influence of several factors on the bolted-joint response. The factors considered included bolt preload, washer-surface-bearing size, edge boundary conditions, joint segment length, and loading history. Joint response is reported in terms of displacements, gap opening, and surface strains. Most of the factors studied were determined to have minimal effect on the bolted-joint response; however, the washer-bearing-surface size affected the response significantly.

Design of wiring for aerospace vehicles relies on an understanding of "ampacity" which refers to the current carrying capacity of wires, either, individually or in wire bundles. Designers rely on standards to derate allowable current flow to prevent exceedance of wire temperature limits due to resistive heat dissipation within the wires or wire bundles. These standards often add considerable margins and are based on empirical data. Commercial providers are taking an aggressive approach to wire sizing which challenges the conventional wisdom of the established standards. Thermal modelling of wire bundles may offer significant mass reduction in a system if the technique can be generalized to produce reliable temperature predictions for arbitrary bundle configurations. Thermal analysis has been applied to the problem of wire bundles wherein any or all of the wires within the bundle may carry current. Wire bundles present analytical challenges because the heat transfer path from conductors internal to the bundle is tortuous, relying on internal radiation and thermal interface conductance to move the heat from within the bundle to the external jacket where it can be carried away by convective and radiative heat transfer. The problem is further complicated by the dependence of wire electrical resistivity on temperature. Reduced heat transfers out of the bundle leads to higher conductor temperatures and, hence, increased resistive heat dissipation. Development of a generalized wire bundle thermal model is presented and compared with test data. The steady state heat balance for a single wire is derived and extended to the bundle configuration. The generalized model includes the effects of temperature varying resistance, internal radiation and thermal interface conductance, external radiation and temperature varying convective relief from the free surface. The sensitivity of the response to uncertainties in key model parameters is explored using Monte Carlo analysis.

A brief history of the X-15-3 adaptive control system and an analysis of the destructive limit-cycle oscillation that occurred during its final November 1967 flight are presented. The X-15 was a piloted single-seat rocket-propelled hypersonic research aircraft operated by the NASA Flight Research Center from 1959 until 1968. Due to the limited information previously available in the public domain and the 1968 decision by the accident investigation board to forego detailed analysis of the adaptive control system’s role in the accident, it was widely assumed in the adaptive controls community that an anomalous behavior of the adaptive component caused the loss of control. Notwithstanding the complex human factors and subsystem failures that contributed to the accident, it is shown that the adaptation dynamics were not a causal factor. The limit cycle observed in the flight data is reproduced in a nonlinear time-domain simulation. Describing function analysis reveals that the instability was caused by a latent design error in the inner-loop structural filters that did not account for the nonlinear behavior of the X-15 servoactuator under rate saturation when coupled with the lightly damped aircraft longitudinal mode at high Mach numbers.

NASA is endeavoring on an ambitious return to the Moon and eventually on to Mars through the Artemis Program leveraging innovative technologies to establish sustainable exploration architectures collaborating with US commercial and international partners [1]. Future NASA architectures have baselined cryogenic propulsion systems to support lunar missions and ultimately future missions to Mars. NASA has been investing in maturing CFM active and passive storage, transfer, and gauging technologies over the last decade plus primarily focused on ground development with a few small-scale microgravity fluid experiments. Recently, NASA created a Cryogenic Fluid Management (CFM) Technology Roadmap identifying the critical gaps requiring further development to reach a technology readiness level (TRL) of 6 prior to infusion to flight applications. To address the technology gaps the Space Technology Mission Directorate (STMD) strategically plans to invest in a diversified CFM portfolio approach through ground and flight demonstrations, collaborating with international partners, and leveraging Public Private Partnerships (PPPs) opportunities with US industry through the Tipping Point and Announcement of Collaborative Opportunities (ACO) solicitations. Once proven, these system capabilities will enable the high performing cryogenic propellant systems needed for the Artemis Program and beyond.

Airbag-based methods for crew impact attenuation have been highlighted as a potential means of easing the mass constraints currently affecting the Orion Crew Exploration Vehicle. An analog airbag test article has been designed, fabricated, and subjected to a drop test campaign to investigate the potential performance of such a system. Through a fusion of accelerometer measurements and photogrammetric analysis of high-speed camera data; it was found that the system performed adequately up to impact velocities of 4.7m/s, where a low-injury risk was maintained throughout the impacting event. This paper presents the background and development behind this first generation system, as well as the results and lessons learnt from this preliminary drop test campaign.

JEANNETTE YEN is the Director of Georgia Tech's Center for Biologically Inspired Design, bringing together biologists, engineers, and physical scientists who seek to

Aerospace and military components must be designed and tested to withstand shock and vibration environments in terms of shock response spectrum (SRS) and power spectral density (PSD) specifications. Shock tests are usually more difficult to configure and control in the test lab. Furthermore, some lower and mid-level SRS specifications may lack the true damage potential to justify shock testing. The purpose of this paper is to demonstrate a comparison method based on the fatigue damage spectrum (FDS) to determine whether the random vibration test covers the shock requirement. This method is found to be effective within a framework of assumptions.

There have been many advancements and accomplishments over the last few years using human modeling for human factors engineering analysis for design of spacecraft. The key methods used for this are motion capture and computer generated human models. The focus of this paper is to explain the human modeling currently used at Kennedy Space Center (KSC), and to explain the future plans for human modeling for future spacecraft designs. <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">1</sup>

Abstract The goals of this two-phase experimental program were to optimize the effectiveness of an icephobic coating for use on several Space Shuttle surfaces. Coating application with a foam brush provided consistent, controlled and reproducible surface coverage. Ice samples were grown slowly and consistently at −10°C prior to cooling to a constant −112°C temperature for cryogenic double lap shear testing. Phase 1 tests were focused on finding an optimal coating mix of Rain-X and varying weight fractions of PTFE powders MP-55 and UF-8TA. The MP-55 coatings produced large reductions in ice adhesion to aluminum coupons while the UF-8TA coatings were similar to uncoated controls. The M4 mixture with 40% MP-55 and 60% Rain-X gave the best and most consistent coating with outstanding performance and durability through five cycles of ice growth and adhesion failure. Phase 2 tests verified the effectiveness and durability of this coating over Koropon, Kapton tape, Kapton film and fire-retardant-paint surfaces on the shuttle and quantified the changes in effectiveness resulting from the addition of an ultraviolet light absorber (UVA). Solvent loss from Rain-X during prolonged mixing of the coating caused a greater increase in ice adhesion than that by adding the UVA.

The NASA Engineering and Safety Center (NESC) is developing the Max Launch Abort System (MLAS) as a risk-mitigation design should problems arise with the baseline Orion spacecraft launch abort design. The Max in MLAS is dedicated to Max Faget, the renowned NASA spacecraft designer. The MLAS flight test vehicle consists of boost skirt, coast skirt and the MLAS fairing which houses a full scale boilerplate Orion Crew Module (CM). The objective of the flight test is to prove that the CM can be released from the MLAS fairing during pad abort conditions without detrimental recontact between the CM and fairing, achieving performance similar to the Orion launch abort system. The boost and coast skirts provide the necessary thrust and stability to achieve the flight test conditions and are released prior to the test -- much like the Little Joe booster was used in the Apollo Launch Escape System tests. To achieve the test objective, two parachutes are deployed from the fairing to reorient the CM/fairing to a heatshield first orientation. The parachutes then provide the force necessary to reduce the total angle of attack and body angular rates required for safe release of the CM from the fairing. A secondary test objective after CM release from the fairing is to investigate the removal of the CM forward bay cover (FBC) with CM drogue parachutes for the purpose of attempting to synchronously deploying a set of CM main parachutes. Although multiple parachute deployments are used in the MLAS flight test vehicle to complete its objective, there are only two parachute types employed in the flight test. Five of the nine parachutes used for MLAS are 27.6 ft D(sub 0) ribbon parachutes, and the remaining four are standard G-12 cargo parachutes. This paper presents an overview of the 27.6 ft D(sub 0) ribbon parachute system employed on the MLAS flight test vehicle for coast skirt separation, fairing reorientation, and as drogue parachutes for the CM after separation from the fairing. Discussion will include: the process used to select this design, previously proven as a spin/stall recovery parachute; descriptions of all components of the parachute system; the minor modifications necessary to adapt the parachute to the MLAS program; the techniques used to analyze the parachute for the multiple roles it performs; a discussion of the rigging techniques used to interface the parachute system to the vehicle; and a brief description of how the evolution of the program affected parachute usage and analysis. An overview of the Objective system, rationale for the MLAS approach and the future of the program will also be presented. We hope to have flight test results to report at the time of the Conference Presentation.

An industry wide survey of GNC sensors, namely star trackers, gyros, and sun sensors was undertaken. Size, mass, power, and various performance metrics were recorded for each category. A multidimensional analysis was performed, looking at the spectrum of available sensors, with the intent of identifying gaps in the available capability range. Mission types that are not currently well served by the available components are discussed, as well as some missions that would be enabled by filling gaps in the component space.

His past experience includes human activity recognition features and surveillance

“Failure is not an option” was a quote from a movie which was incorrectly attributed to Apollo 13 Flight Director, Gene Kranz. Although this phrase was popularized and made famous by the movie “Apollo 13”, it was never actually uttered by Gene Kranz. While never spoken, this philosophy was, however, deeply engrained in the culture of the Mission Operations Directorate (MOD) at the Johnson Space Center (JSC) which Mr. Kranz was instrumental in developing. While this philosophy/mantra might suit an “operational” vehicle which supports human life, it can have disastrous consequences when applied to operation and/or development of experimental or research vehicles. In fact, the two Space Shuttle tragedies resulted, in part, from organizational behaviors which relied on a “can-do” spirit and a professed belief that the Space Shuttle was an operational vehicle. Following NASA’s most recent tragedy, the Columbia accident, NASA adopted a probabilistic analysis of risk as a major element in its equation for assessing launch readiness during subsequent Flight Readiness Reviews (FRR). This paper will discuss the misuse of probability calculations to develop flight rationale at key FRR reviews and, hence, questions the use of such methodology for the selection and design of next-generation launch vehicles, where much is uncertain and sensitivities of these uncertainties are multiplied.

The first flight of NASA's new exploration-class launch vehicle, the Space Launch System (SLS), will test a myriad of systems designed to enable the next generation of deep space human spaceflight, and launch from Kennedy Space Center no earlier than December 2019. The initial Block 1 configuration for Exploration Mission 1 (EM-1)will be capable of lofting at least 70 metric tons (t)of payload and send the Orion crew vehicle into a distant retrograde lunar orbit, paving the way for future crew missions to cislunar space and eventually Mars. A Block 1B version of SLS will lift at least 34 t to trans-lunar injection (TLI)in its crew configuration and at least 37 t to TLI in its cargo configuration no earlier than 2024. For Mars-class payloads, larger fairings and payload adapters for the Block 2 cargo vehicle are under consideration. For missions beyond the Earth-Moon system, SLS offers greater characteristic energy (C3)than any other launch vehicle, enabling shorter transit times or heavier payloads with more robust science packages for missions to the outer solar system. Indeed, the unmatched combination of thrust, payload volume and departure energy that SLS provides opens new opportunities for human and robotic exploration of deep space. To support the delivery of infrastructure on all of these flights, a family of SLS Payload Adapters (PLA)is being developed to provide ELV class (1575-mm, 2624-mm, 4394-mm)and larger spacecraft/payload interfaces for both crewed (Orion)and cargo (fairing)missions. These PLAs also provide the potential of accommodating various configurations of 6U, 12U and 27U Secondary Payloads (SPL). Work on demonstrating the manufacturing of these 8.4-m diameter composite structures is already in progress at Marshall Space Flight Center in Huntsville, Alabama, which manages the SLS Program. Because of the many potential configurations required to support SLS missions ranging from sending Europa Clipper to Jovian space to establishing a lunar orbiting Gateway, there is a critical need for establishing the fewest PLA designs that can accommodate the most SLS payloads possible. This paper will summarize applications from a NASA Engineering and Safety Center (NESC)led Model Based Systems Engineering (MBSE)pathfinder activity to develop a “digital” PLA feasibility assessment approach. This approach will help potential users optimize their interface to SLS by providing analysts with the means to reduce PLA feasibility definition cycle time/effort by over 75%. This also allows more feasibility assessment “turns” available to single and multiple payload elements on a single SLS launch. This translates into providing users with options that allows them to optimize upmass available to payload versus being required for PLA structure.

Long duration human space exploration demands unprecedented levels of automation, to carry the load of vehicle management for the new explorers, and assist their exploration work with new capability. Automated and robotic systems are probably sophisticated and sturdy enough to do this work -- but such systems have never been human-rated like all other NASA systems used in human space flight. We present here a perspective on architecture and requirements for the interfaces and interactions between human explorers and their array of automated systems; and we present a necessary approach to human-rate the systems for the space program: for surface operations as well as for in-flight monitoring and control. We continue to hope this topic will be an invitation to dialog and to consideration of a difficult issue that will face new generations of explorers and their supporters back on Earth.