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NASA NTRSDate not supplied

VERNE: Revealing the Mysteries and Histories of Venus

Introduction: The three Venus missions that were recently selected for upcoming flight (VERITAS, DAVINCI+, and EnVision) will be incredibly valuable to our understanding of Venus’ history, geology, and atmosphere. However, even once completed, key gaps in our knowledge of Venus, and more generally the formation and active processes on rocky, Earth-like planets, will still persist. Remaining questions include 1) how global intrinsic magnetic fields might be maintained on rocky worlds, and how they could then go extinct, and 2) what role atmospheric sulfur chemistry plays in climates of Earth-like planets, which is an increasingly timely subject as Earth’s own atmospheric sulfur content is climbing due to human activity. These questions require in-situ observations from Venus’ cloud deck, at the altitudes at which the UV absorber exists. The Venus Environment Research and Novel Exploration (VERNE) mission will address these questions with an aerial platform that will drift around the equatorial region of the planet for 9 days. VERNE will collect data to determine the identity of the mysterious UV absorber, while also taking magnetic field measurements over the tesserae, the regions on Venus that are most likely to retain remanent crustal magnetization, in order to understand the potential role of a past intrinsically-generated global magnetic field on Venus. Mission Objectives: The two major science objectives that drive the VERNE mission are: 1) Determine the identity of the Venusian unknown ultraviolet absorber(s). First observed approximately a century ago [1], the composition of Venus’ ultraviolet (UV) absorber is one of the oldest mysteries in Venus atmospheric chemistry [2,3]. While several candidate UV absorbers (mostly sulfur species) have been proposed, no consensus has been reached on its composition and its specific interactions with the atmosphere. Determining the identity of the UV absorber will aid climate models by showing how and where incident solar energy is absorbed by the atmosphere and make progress towards understanding the chemical and energetic processes taking place above Venus’ upper cloud deck [4]. 2) Determine if Venus retains evidence of a past, internally-generated magnetic field. While Venus does not currently have an intrinsically-generated magnetic field, evidence for the existence of a past field on Venus and a timeline of its decay will fill in a more holistic picture of the evolution of Venus’ geological record and atmosphere. As the oldest geologic units on the surface, Venus’ tesserae may still have remanent crustal magnetization signatures within the rocky composition [5]. In any case, the signatures detected will provide insight into Venus’ past geological and core dynamo activity. Mission Summary: VERNE includes a 3-part flight system made up of 1) an entry system with a HEEET (Heatshield for Extreme Entry Environment Technology) aeroshell, 2) an orbiter for data relay, and 3) an in-situ balloon and gondola. After entry into Venus’ atmosphere, the balloon will be deployed within the upper cloud deck, at an altitude of 62 km, over a tesserae region. The in-situ data collection will last for the duration of 2 full circumnavigations of the planet, which will take ~9 days. Instrumentation Suite: The four instruments that comprise VERNE’s instrument payload will enable the identification of the unknown UV absorber and the detection of remanent crustal magnetization if it exists in the tesserae. The proposed instrument suite cycle during mission operations is shown in Fig. 1. The four instruments are described below: 1) Adams (ion neutral mass spectrometer) has a range of 18-257 AMU and will detect and distinguish the mixing ratios of O₂, H₂O, H₂SO₄, S, S₂, S₃, S₄, S₅, S₆, S₇, S₈, SO, SO₂, OSSO, SO₃, Cl₂, FeCl₃ and other trace sulfur and organic species. The spatial (longitudinal) and temporal (day/night) variations will be observed throughout 2 circumnavigations with a sample cadence of 12 minutes. 2) Shelley (nephelometer) will determine the size distribution of the aerosols (0.4 to 36 um) in the atmosphere. With a size resolution of <0.7 um, it can determine which mode of H2SO4 is present and characterize the large (>30 um) organic particles previously detected by the Venera and Galileo missions [6,7]. 3) Herbert (UV imager) will measure UV radiance at 283 nm (the wavelength of SO2 absorption) and 365 nm (the unknown part of the absorber). UV images will be taken concurrently with the INMS and nephelometer to correlate UV absorption with abundances of the UV absorber species. 4) Vonnegut (magnetometer) has a range of >600 nT and a precision and accuracy of 1 nT. If magnetized by a past field, the crust may have retained a magnetization of up to 3 A/m2 [5]. With a noise floor of 10 nT, the magnetometer will be able to detect RCM from an altitude of 62 km, even if the thickness of the magnetized crust is just 1 km (Fig. 2). Mission Concept Design: VERNE will be launched with a mass of 3300 kg in an intermediate-high performance class vehicle with a 4-m fairing. The 475-day mission includes 466 days for the cruise, coasting, and orbit initialization phases before entry, descent, and balloon deployment. During the 9-day science phase, the aerial platform will make 2 circumnavigations of the planet at an altitude of 62 km. The INMS and the nephelometer will acquire data for 2 hours during the daytime and nighttime during each circumnavigation, while the UV imager will be on for the duration of the daytime, and the magnetometer will be operational throughout the entirety of the science phase. Data will be stored and processed with the JPL-designed Sphinx command and data handling system. Data will be sent from the balloon to the orbiter using an S-band relay link, stored on the orbiter, and then forwarded to Earth where it will be received by the DSN. Conclusion: VERNE will fill key gaps in our understanding of the history and ongoing processes related to the geology and atmosphere of Venus and rocky worlds in general. Even with adequate flight system contingencies and expected costs below the $900M New Frontiers cost cap, VERNE is not without its risks and challenges. Further trade spaces to explore include 1) using solely battery power vs. including solar panels to increase the mission duration and 2) investigating the use of lightweight materials and 3D-printed structures to reduce the gondola mass, among others. Acknowledgments: We would like to thank the JPL Planetary Science Summer School, especially our mentors Troy Hudson, Karl Mitchell, and Leslie Lowes, as well as our Team-X study lead Al Nash and the members of Team-X. We’d additionally like to thank our review panel for asking insightful questions and providing valuable feedback. References: [1] Ross, F. E. (1928) Astrophysical J., 68, 57-92. [2] Rossow, W. B. et al. (1980) J. Geophysical Research, 85, 8107-8128. [3] Pinto, J. P. et al. (2021) Nature Communications, 12, 175. [4] Titov, D. V. et al. (2007) Cosmic Research, 21, 401. [5] O’Rourke, J. et al. (2019) Geophysical Research Lett., 46, 5768–5777. [6] Limaye, S. S. et al. (2018) Astrobiology, 18(9), 1181-1198. [7] Grinspoon, D. H. et al. (2013) Planetary and Space Sci., 41(7), 515-542. [8] Parker, R. L. (2003) J. Geophysical Research, 108, 5006. *The cost information contained in this document is of a budgetary and planning nature and is intended for informational purposes only. It does not constitute a commitment on the part of JPL and/or Caltech.

H Alpert↗
NASA NTRS2021

NASA SpaceCube Edge TPU SmallSat Card for Autonomous Operations and Onboard Science-Data Analysis

Using state-of-the-art artificial intelligence (AI)frameworks onboard spacecraft is challenging because common spacecraft processors cannot provide comparable performance to datacenters with server-grade CPUs and GPUs available for terrestrial applications and advanced deep-learning networks. This limitation makes small, lo w-p o we r AI microchip architectures, such as the Google Coral Edge Tensor Processing Unit (TPU), attractive for space missions where the application-specific design enables both high-performance and power-efficient computing for AI applications. To address these challenging considerations for space deployment, this research introduces the design and capabilities of a CubeSat-sized Edge TPU-based co-processor card, known as the SpaceCube Low-power Ed g e Artificial Intelligence Resilient Node (SC-LEARN). This design conforms to NASA’s CubeSat Card Specification (CS2) for integration into next-generation SmallSat and CubeSat systems. This paper describes the overarching architecture and design of the SC-LEARN, as well as, the supporting test card designed for rapid prototyping and evaluation. The SC-LEARN was developed with three operational modes: (1) a high-performance parallel-processing mode,(2)a fault-tolerant mode for onboard resilience, and (3) a power-saving mode with cold spares. Importantly, this research also elaborates on both training and quantization of Tensor Flow models for the SC-LEARN for use onboard with representative, open-source datasets. Lastly, we describe future research plans, including radiation-beam testing and flight demonstration.

Advanced avionics↗
NASA NTRS1972

Development of termination and utilization concepts for flat conductor cables. Volume 3: Cost study comparison, flat versus round conductor cable

A cost study comparing flat conductor cable (FCC) with small-gage wire (SGW) and conventional round conductor cable (RCC) is presented. This study was based on a vehicle wiring system consisting of 110,000 ft of conventional RCC equally divided between AWG sizes 20,22, and 24 using MIL-W-81044-type wire and MIL-C-26500 circular connectors. Basic cost data were developed on a similar-sized commercial jet airplane wiring system on a previous company R&D program in which advanced wiring techniques were carried through equivalent installations on an airplane mockup; and on data developed on typical average bundles during this program. Various cost elements included were engineering labor, operations (manufacturing) labor, material costs, and cost impact on payload. Engineering labor includes design, wiring system integration, wiring diagrams and cable assembly drawings, wire installations, and other related supporting functions such as the electronic data processing for the wiring. Operations labor includes mockup, tooling and production planning, fabrication, assembly, installation, and quality control cost impact on payload is the conversion of wiring system weight variations through use of different wiring concepts to program payload benefits in terms of dollars.

Source record↗
DOE OSTI2025

SPC-71260 Rev 0 MARVEL Heat Extraction Subsystem Secondary Coolant Equipment (SCE) Design/Build

A. The Microreactor Applications Research Validation and Evaluation (MARVEL) reactor will offer experimental capabilities that are not currently available at DOE’s national laboratories. Idaho National Laboratory (INL), operated for the U.S. Department of Energy (DOE) by Battelle Energy Alliance, LLC (BEA) (Contractor hereafter) is procuring services for the design, analysis, fabrication, testing and delivery of a Secondary Coolant Equipment system (SCE). This specification contains the requirements for design, analysis, fabrication, testing and delivery of the SCE as described herein. The MARVEL reactor is a microreactor which uses eutectic sodium-potassium alloy (NaK) as a primary coolant. The primary coolant is circulated through four primary loops by natural convection of the coolant. In each loop is a closed well which will accommodate an intermediate heat exchanger (IHX) for extracting heat from the loop. These wells will be referred to in this specification as the “IHX wells.” It is intended for the IHX containment to also be filled with NaK. The MARVEL design team has determined that a Heat Extraction System (HES) using pumped NaK will be used to extract heat from the IHXs and deliver it to a downstream system for power generation or alternate process heat users. This Heat Extraction System will enable MARVEL operations including the ability to test, demonstrate, and address issues related to installation, startup, and operations. In addition, it will allow down-stream utilization of process heat for various uses. The objective of this specification is to develop the final design for the HES Secondary Coolant Equipment system (SCE) that will be used as the core of the HES. This system provides control of the NaK circulation between the MARVEL reactor and the subsequent process heat utilization systems. It does not include design of the Intermediate Heat Exchangers and piping inside the T-REXc pit in which the reactor is located. B. The MARVEL microreactor will be installed in the Transient Reactor Test Facility (TREAT) building in the Transient Reactor Test (TREAT) Micro-Reactor Experiment Cell (T-REXc) C. An INL Subcontractor has developed a conceptual design for this system per SPC-71145, referred to in that specification as the Process Heat Extraction System. SPC-71260 is based on the pumped NaK loop concept developed under SPC-71145. D. The SCE system design and (as option scope) fabrication shall be provided by the awardee of the subcontract (Subcontractor hereafter) pertaining to this Specification. Prior to shipment, the SCE will be fabricated, assembled, and tested at the Subcontractor’s facility. After successful completion of acceptance testing, the SCE and associated equipment will be shipped to the Materials and Fuels Complex (MFC) at the INL (Contractor’s Facility hereafter) to be installed by others in TREAT/T-REXc.

21 - SPECIFIC NUCLEAR REACTORS AND ASSOCIATED PLAN↗
NASA NTRS2011

Full-Scaled Advanced Systems Testbed: Ensuring Success of Adaptive Control Research Through Project Lifecycle Risk Mitigation

The National Aeronautics and Space Administration's Dryden Flight Research Center completed flight testing of adaptive controls research on the Full-Scale Advance Systems Testbed (FAST) in January of 2011. The research addressed technical challenges involved with reducing risk in an increasingly complex and dynamic national airspace. Specific challenges lie with the development of validated, multidisciplinary, integrated aircraft control design tools and techniques to enable safe flight in the presence of adverse conditions such as structural damage, control surface failures, or aerodynamic upsets. The testbed is an F-18 aircraft serving as a full-scale vehicle to test and validate adaptive flight control research and lends a significant confidence to the development, maturation, and acceptance process of incorporating adaptive control laws into follow-on research and the operational environment. The experimental systems integrated into FAST were designed to allow for flexible yet safe flight test evaluation and validation of modern adaptive control technologies and revolve around two major hardware upgrades: the modification of Production Support Flight Control Computers (PSFCC) and integration of two, fourth-generation Airborne Research Test Systems (ARTS). Post-hardware integration verification and validation provided the foundation for safe flight test of Nonlinear Dynamic Inversion and Model Reference Aircraft Control adaptive control law experiments. To ensure success of flight in terms of cost, schedule, and test results, emphasis on risk management was incorporated into early stages of design and flight test planning and continued through the execution of each flight test mission. Specific consideration was made to incorporate safety features within the hardware and software to alleviate user demands as well as into test processes and training to reduce human factor impacts to safe and successful flight test. This paper describes the research configuration, experiment functionality, overall risk mitigation, flight test approach and results, and lessons learned of adaptive controls research of the Full-Scale Advanced Systems Testbed.

Pavlock, Kate M.↗
NASA NTRS2010

Public Risk Assessment Program

The Public Entry Risk Assessment (PERA) program addresses risk to the public from shuttle or other spacecraft re-entry trajectories. Managing public risk to acceptable levels is a major component of safe spacecraft operation. PERA is given scenario inputs of vehicle trajectory, probability of failure along that trajectory, the resulting debris characteristics, and field size and distribution, and returns risk metrics that quantify the individual and collective risk posed by that scenario. Due to the large volume of data required to perform such a risk analysis, PERA was designed to streamline the analysis process by using innovative mathematical analysis of the risk assessment equations. Real-time analysis in the event of a shuttle contingency operation, such as damage to the Orbiter, is possible because PERA allows for a change to the probability of failure models, therefore providing a much quicker estimation of public risk. PERA also provides the ability to generate movie files showing how the entry risk changes as the entry develops. PERA was designed to streamline the computation of the enormous amounts of data needed for this type of risk assessment by using an average distribution of debris on the ground, rather than pinpointing the impact point of every piece of debris. This has reduced the amount of computational time significantly without reducing the accuracy of the results. PERA was written in MATLAB; a compiled version can run from a DOS or UNIX prompt.

Mendeck, Gavin↗
NASA NTRS2001

Lightweight Chambers for Thrust Assemblies

The Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) has successfully applied new materials and fabrication techniques to create actively cooled thrust chambers that operate 200-400 degrees hotter and weigh 50% lighter than conventional designs. In some vehicles, thrust assemblies account for as much as 20% of the engine weight. So, reducing the weight of these components and increasing their operating range will benefit many engines and vehicle designs, including Reusable Launch Vehicle (RLV) concepts. Obviously, copper and steel alloys have been used successfully for many years in the chamber components of thrust assemblies. Yet, by replacing the steel alloys with Polymer Matrix Composite (PMC) and/or Metal Matrix Composite (MMC) materials, design weights can be drastically reduced. In addition, replacing the traditional copper alloys with a Ceramic Matrix Composite (CMC) or an advanced copper alloy (Cu-8Cr-4Nb, also known as GRCop-84) significantly increases allowable operating temperatures. Several small MMC and PMC demonstration chambers have recently been fabricated with promising results. Each of these designs included GRCop-84 for the cooled chamber liner. These units successfully verified that designs over 50% lighter are feasible. New fabrication processes, including advanced casting technology and a low cost vacuum plasma spray (VPS) process, were also demonstrated with these units. Hot-fire testing at MSFC is currently being conducted on the chambers to verify increased operating temperatures available with the GRCop-84 liner. Unique CMC chamber liners were also successfully fabricated and prepared for hot-fire testing. Yet, early results indicate these CMC liners need significantly more development in order to use them in required chamber designs. Based on the successful efforts with the MMC and PMC concepts, two full size "lightweight" chambers are currently being designed and fabricated for hot-fire testing at MSFC in 2001. These "full size" chambers will be similar in size to those used on the X33 engine (RS2200). One will be fabricated with a MMC structural jacket, while the other uses a PMC jacket. Each will be designed for thrust levels of 15,000 pounds in an oxygen/hydrogen environment with liquid hydrogen coolant. Both chambers will use GRCop-84 for its channel wall liner. Each unit is expected to be at least 60% lighter than a conventional design with traditional materials. Hot-fire testing on the full size units in late 2001 will directly compare performance results between a conventional chamber design and these "lightweight" alternatives. The technology developed and demonstrated by this effort will not only benefit next generation RLV programs, but it can be applied to other existing and future engine programs, as well. Efforts were sponsored by the Advanced Space Transportation Program for RLV Focused Technologies. The task team was led by MSFC with additional members from NASA-Glenn Research Center and the Rocketdyne Division of The Boeing Company. Specific materials development and fabrication processes were provided by Aerojet, Lockheed Martin Astronautics, Composite Optics, Inc., Hyper-Therm, Ceramic Composites, Inc., MSE Technology Applications, and Plasma Processes, Inc.

Elam, Sandra K.↗
NASA NTRS1994

Microfabricated Hydrogen Sensor Technology for Aerospace and Commercial Applications

Leaks on the Space Shuttle while on the Launch Pad have generated interest in hydrogen leak monitoring technology. An effective leak monitoring system requires reliable hydrogen sensors, hardware, and software to monitor the sensors. The system should process the sensor outputs and provide real-time leak monitoring information to the operator. This paper discusses the progress in developing such a complete leak monitoring system. Advanced microfabricated hydrogen sensors are being fabricated at Case Western Reserve University (CWRU) and tested at NASA Lewis Research Center (LeRC) and Gencorp Aerojet (Aerojet). Changes in the hydrogen concentrations are detected using a PdAg on silicon Schottky diode structure. Sensor temperature control is achieved with a temperature sensor and heater fabricated onto the sensor chip. Results of the characterization of these sensors are presented. These sensors can detect low concentrations of hydrogen in inert environments with high sensitivity and quick response time. Aerojet is developing the hardware and software for a multipoint leak monitoring system designed to provide leak source and magnitude information in real time. The monitoring system processes data from the hydrogen sensors and presents the operator with a visual indication of the leak location and magnitude. Work has commenced on integrating the NASA LeRC-CWRU hydrogen sensors with the Aerojet designed monitoring system. Although the leak monitoring system was designed for hydrogen propulsion systems, the possible applications of this monitoring system are wide ranged. Possible commercialization of the system will also be discussed.

Hunter, Gary W.↗
NASA NTRS2017

SpaceCubeX: A Framework for Evaluating Hybrid Multi-Core CPU FPGA DSP Architectures

The SpaceCubeX project is motivated by the need for high performance, modular, and scalable on-board processing to help scientists answer critical 21st century questions about global climate change, air quality, ocean health, and ecosystem dynamics, while adding new capabilities such as low-latency data products for extreme event warnings. These goals translate into on-board processing throughput requirements that are on the order of 100-1,000 more than those of previous Earth Science missions for standard processing, compression, storage, and downlink operations. To study possible future architectures to achieve these performance requirements, the SpaceCubeX project provides an evolvable testbed and framework that enables a focused design space exploration of candidate hybrid CPU/FPGA/DSP processing architectures. The framework includes ArchGen, an architecture generator tool populated with candidate architecture components, performance models, and IP cores, that allows an end user to specify the type, number, and connectivity of a hybrid architecture. The framework requires minimal extensions to integrate new processors, such as the anticipated High Performance Spaceflight Computer (HPSC), reducing time to initiate benchmarking by months. To evaluate the framework, we leverage a wide suite of high performance embedded computing benchmarks and Earth science scenarios to ensure robust architecture characterization. We report on our projects Year 1 efforts and demonstrate the capabilities across four simulation testbed models, a baseline SpaceCube 2.0 system, a dual ARM A9 processor system, a hybrid quad ARM A53 and FPGA system, and a hybrid quad ARM A53 and DSP system.

Hybrid Flight Architectures↗
NASA NTRS1990

Experiment design and operations

The objectives, design, and field operations of the First ISLSCP Field Experiment (FIFE) are described. The simultaneous acquisition of satellite, atmosphere, and surface data, and the understanding of the processes governing surface energy and mass exchange and how these are manifested in satellite-resolution radiometric data are identified as the specific objectives of the field-phase experiment. The central issues concerning the design of the field experiment are considered: the size of the site, the duration of the experiment, and the location of the site; it is noted that the Konza Prairie National Reserve was selected as the focus of the study. Field operations in 1987 and 1989 are discussed, and it is pointed out that a data set is available now from a single combined repository to all FIFE investigators, and that scientists can test models and algorithms on scales consistent with satellite observations and with enough supporting data on finer scales.

Sellers, P. J.↗
NASA NTRS1992

A VLSI implementation of DCT using pass transistor technology

A VLSI design for performing the Discrete Cosine Transform (DCT) operation on image blocks of size 16 x 16 in a real time fashion operating at 34 MHz (worst case) is presented. The process used was Hewlett-Packard's CMOS26--A 3 metal CMOS process with a minimum feature size of 0.75 micron. The design is based on Multiply-Accumulate (MAC) cells which make use of a modified Booth recoding algorithm for performing multiplication. The design of these cells is straight forward, and the layouts are regular with no complex routing. Two versions of these MAC cells were designed and their layouts completed. Both versions were simulated using SPICE to estimate their performance. One version is slightly faster at the cost of larger silicon area and higher power consumption. An improvement in speed of almost 20 percent is achieved after several iterations of simulation and re-sizing.

Kamath, S.↗
NASA NTRS1984

Technical Workshop: Advanced Helicopter Cockpit Design

Information processing demands on both civilian and military aircrews have increased enormously as rotorcraft have come to be used for adverse weather, day/night, and remote area missions. Applied psychology, engineering, or operational research for future helicopter cockpit design criteria were identified. Three areas were addressed: (1) operational requirements, (2) advanced avionics, and (3) man-system integration.

Hemingway, J. C.↗
NASA NTRS2017

Evaluation of a Candidate Trace Contaminant Control Subsystem Architecture: The High Velocity, Low Aspect Ratio (HVLA) Adsorption Process

Traditional gas-phase trace contaminant control adsorption process flow is constrained as required to maintain high contaminant single-pass adsorption efficiency. Specifically, the bed superficial velocity is controlled to limit the adsorption mass-transfer zone length relative to the physical adsorption bed; this is aided by traditional high-aspect ratio bed design. Through operation in this manner, most contaminants, including those with relatively high potential energy are readily adsorbed. A consequence of this operational approach, however, is a limited available operational flow margin. By considering a paradigm shift in adsorption architecture design and operations, in which flows of high superficial velocity are treated by low-aspect ratio sorbent beds, the range of well-adsorbed contaminants becomes limited, but the process flow is increased such that contaminant leaks or emerging contaminants of interest may be effectively controlled. To this end, the high velocity, low aspect ratio (HVLA) adsorption process architecture was demonstrated against a trace contaminant load representative of the International Space Station atmosphere. Two HVLA concept packaging designs (linear flow and radial flow) were tested. The performance of each design was evaluated and compared against computer simulation. Utilizing the HVLA process, long and sustained control of heavy organic contaminants was demonstrated.

Kayatin, Matthew J.↗
NASA NTRS2022

Mars 2020 Entry, Descent, and Landing Software Implementation

On February 18th, 2021, the Mars 2020 project's Perseverance Rover successfully touched down on the Martian surface after nearly eight years of development. The Mars 2020 Entry, Descent, and Landing (EDL) System largely leveraged heritage from the Mars Science Laboratory (MSL) EDL System while employing targeted technological advancements. The landing process is autonomously directed by a software behavior implemented in the rover's primary flight computer called the EDL Timeline that assumes control of the vehicle six days before atmospheric entry. In addition to performing the critical function of landing the rover on the Martian surface, the EDL timeline behavior must co-exist in a non-partitioned software and system environment with other high-level functions that accomplish the goals for the rest of the mission. Due to the criticality of EDL, the potential for loss of mission, and a need for complete system autonomy, the standard for how the EDL Timeline interacts with other functions in the system is highly constrained. This paper first walks through the basics of the EDL Timeline mechanics and how the behavior is designed to account for internal system variations and environmental unknowns. It then summarizes the interactions between the EDL timeline and other high-level system behaviors like spacecraft mode transitions and system fault protection, focusing on the complications that arise when passing spacecraft control between executive functions. Although the MSL-inherited EDL System is reliable and capable, targeted updates and a thorough verification and validation program were required for Mars 2020. This paper discusses changes made to close vulnerabilities discovered during both MSL and Mars 2020 development cycles, landing system capability enhancements that were enabling for Mars 2020's mission, and how these updates were integrated with the heritage system. It then describes how both analysis and testing campaigns were utilized to verify and validate all aspects of EDL and system behaviors that run during the six days before landing, as well as the operational workarounds that were needed to address problems found during the development and commissioning process. Finally, this paper imparts lessons learned from Mars 2020 EDL development, implementation, and operations, emphasizing how systems designed to conduct time-critical mission events with low margin of error can be improved in the future.

Stehura, Aaron↗
NASA NTRS1979

Electrochemically regenerable carbon dioxide absorber

Preliminary designs were generated for two electrochemically regenerable carbon dioxide absorber concepts. Initially, an electrochemically regenerable absorption bed concept was designed. This concept incorporated the required electrochemical regeneration components in the absorber design, permitting the absorbent to be regenerated within the absorption bed. This hardware was identified as the electrochemical absorber hardware. The second hardware concept separated the functional components of the regeneration and absorption process. This design approach minimized the extravehicular activity component volume by eliminating regeneration hardware components within the absorber. The electrochemical absorber hardware was extensively characterized for major operating parameters such as inlet carbon dioxide partial pressure, process air flow rate, operational pressure, inlet relative humidity, regeneration current density and absorption/regeneration cycle endurance testing.

Woods, R. R.↗
NASA NTRS2001

Stage Separation Performance Analysis Project

Stage separation process is an important phenomenon in multi-stage launch vehicle operation. The transient flowfield coupled with the multi-body systems is a challenging problem in design analysis. The thermodynamics environment with burning propellants during the upper-stage engine start in the separation processes adds to the complexity of the-entire system. Understanding the underlying flow physics and vehicle dynamics during stage separation is required in designing a multi-stage launch vehicle with good flight performance. A computational fluid dynamics model with the capability to coupling transient multi-body dynamics systems will be a useful tool for simulating the effects of transient flowfield, plume/jet heating and vehicle dynamics. A computational model using generalize mesh system will be used as the basis of this development. The multi-body dynamics system will be solved, by integrating a system of six-degree-of-freedom equations of motion with high accuracy. Multi-body mesh system and their interactions will be modeled using parallel computing algorithms. Adaptive mesh refinement method will also be employed to enhance solution accuracy in the transient process.

Chen, Yen-Sen↗