ALBERT

Description: Launched June 18, 2009 on an Atlas V rocket, NASA's Lunar Reconnaissance Orbiter (LRO) is the first step in NASA's Vision for Space Exploration program and for a human return to the Moon. The spacecraft (SC) carries a wide variety of scientific instruments and provides an extraordinary opportunity to study the lunar landscape at resolutions and over time scales never achieved before. The spacecraft systems are designed to enable achievement of LRO's mission requirements. To that end, LRO's mechanical system employed two two-axis gimbal assemblies used to drive the deployment and articulation of the Solar Array System (SAS) and the High Gain Antenna System (HGAS). This paper describes the design, development, integration, and testing of Gimbal Control Electronics (GCE) and Actuators for both the HGAS and SAS systems, as well as flight testing during the on-orbit commissioning phase and lessons learned.

Description: For more than a decade, several teams have assessed designs for a long-duration free-space human habitat beyond low-Earth orbit (LEO), building upon years of hard-won experience with the International Space Station (ISS). These systems would enable multiple achievements for science and human space flight. Most were intended to be deployed using available or near-future capabilities within about a decade after funding begins and serve as the first major human "stepping stone" beyond LEO. Last year, Thronson and Talay summarized work up to that time on expandable or inflatable concepts for deployment at an Earth-Moon (E-M) L1 or L2 location. Here we summarize our team's more recent work both on a long-duration human habitat that could be deployed beyond LEO within a decade and on the priority goals that such a habitat might accomplish. Particulars of this and other concepts for human operations in cis-lunar space are posted on the web and will be presented at professional conferences, and detailed in future publications by our group.

Description: The importance of accurately pointing spacecraft to our daily lives is pervasive, yet somehow escapes the notice of most people. In this section, we will summarize the processes and technologies used in designing and operating spacecraft pointing (i.e. attitude) systems.

Description: The Lunar Reconnaissance Orbiter (LRO), a spacecraft designed and built at the National Aeronautics and Space Administration s (NASA) Goddard Space Flight Center (GSFC) in Greenbelt, MD, was launched on June 18, 2009 from Cape Canaveral. It is currently in orbit about the Moon taking detailed science measurements and providing a highly accurate mapping of the suface in preparation for the future return of astronauts to a permanent moon base. Onboard the spacecraft is a complex set of algorithms designed by the attitude control engineers at GSFC to control the pointig for all operational events, including anomalies that require the spacecraft to be put into a well known attitude configuration for a sufficiently long duration to allow for the investigation and correction of the anomaly. GSFC level requirements state that each spacecraft s control system design must include a configuration for this pointing and lso be able to maintain a thermally safe and power positive attitude. This stable control algorithm for anomalous events is commonly referred to as the safe mode and consists of control logic thatwill put the spacecraft in this safe configuration defined by the spacecraft s hardware, power and environment capabilities and limitations. The LRO Sun Safe mode consists of a coarse sun-pointing set of algorithms that puts the spacecraft into this thermally safe and power positive attitude and can be achieved wihin a required amount of time from any initial attitude, provided that the system momentum is within the momentum capability of the reaction wheels. On LRO the Sun Safe mode makes use of coarse sun sensors (CSS), an inertial reference unit (IRU) and reaction wheels (RW) to slew the spacecraft to a solar inertial pointing. The CSS and reaction wheels have some level of redundancy because of their numbers. However, the IRU is a single-point-failure piece of hardware. Without the rate information provided by the IRU, the Sun Safe control algorithms could not maintain the required pointing, so a sub-mode of the Sun Safe mode that does not use the IRU was designed. This submode, referred to as the Sun Safe Gyroless control mode, consists of an algorithm that estimates rate information from the CSS and the RW measurements. RW momentum information is used to estimate the body rate parallel to the target sunline, which CSS alone would not be able to observe. Sun Safe can be autonomously, or via ground command, entered from any other control mode and in the event the IRU is not providing rate information, the control mode is switched to the gyroless submode. This paper looks at the design of the Sun Safe modes and discusses the constraints placed on the algorithm and how the mode wored around these constraints. Items of particular interest include CSS placement on the Solar Array (SA) and its implications to design, estimation of body rate information for the Sun Safe Gyroless control mode, and the effect of solar eclipse on each of the Sun Safe modes. Placing CSS on the SA was necessary for the means to put the Sun along the targeted sun-line, nominally normal to the SA panels, for all operational considerations. This had design implications for determining a sun vector during normal SA operations, if one or both gimbals become inoperable and when the SA is in a stowed configuration. The ability of body rate estimation in Sun Safe Gyroless not only uses CSS sun vector data but requires RW momentum measuremens to estimate rates parallel to the sun-line. LRO encounters solar eclipses of some length for most of its orbits about the Moon. With the lack of CSS measurement data a design was implemented in both Sun Safe and Sun Safe Gyroless, they differ because of having or not having IRU measurement data, to carry the spacecraft through these eclipse periods. This paper also includes some discussion of sun avoidance and how it affected design decisions during nominal and eclipse perids for each of the Sun Safe modes.

Description: A hypersonic flowfield model that treats electronic levels of the dominant afterbody radiator, N, as individual species is presented. This model allows electron-ion recombination rate and two-temperature modeling improvements, the latter which are shown to decrease afterbody radiative heating by up to 30%. This increase is primarily due to the addition of the electron-impact-excitation energy-exchange term to the energy equation governing the vibrational-electronic-electron temperature. This model also allows the validity of the often applied quasi-steady state (QSS) approximation to be assessed. The QSS approximation is shown to fail throughout most of the afterbody region for lower electronic states, although this impacts the radiative intensity reaching the surface by less than 15%. By computing the electronic state populations of N within the flowfield solver, instead of through the QSS approximation in the radiation solver, the coupling of nonlocal radiative transition rates to the species continuity equations becomes feasible. Implementation of this higher- fidelity level of coupling between the flowfield and radiation solvers is shown to increase the afterbody radiation by up to 50% relative to the conventional model.

Description: Hypervelocity impacts were performed on six unstressed and six stressed titanium coupons with aluminium: shielding in order to assess the effects of the partial penetration damage on the post impact micromechanical properties of titanium and on the residual strength after impact. This work is performed in support of the defInition of the penetration criteria of the propellant and oxidizer tanks dome surfaces for the service module of the crew exploration vehicle where such a criterion is based on testing and analyses rather than on historical precedence. The objective of this work is to assess the effects of applied biaxial stress on the damage dynamics and morphology. The crater statistics revealed minute differences between stressed and unstressed coupon damage. The post impact residual stress analyses showed that the titanium strength properties were generally unchanged for the unstressed coupons when compared with undamaged titanium. However, high localized strains were shown near the craters during the tensile tests.

Description: The "Stardust" heat shield, composed of a PICA (Phenolic Impregnated Carbon Ablator) Thermal Protection System (TPS), bonded to a composite aeroshell, contains important features which chronicle its time in space as well as re-entry. To guide the further study of the Stardust heat shield, NASA reviewed a number of techniques for inspection of the article. The goals of the inspection were: 1) to establish the material characteristics of the shield and shield components, 2) record the dimensions of shield components and assembly as compared with the pre-flight condition, 3) provide flight infonnation for validation and verification of the FIAT ablation code and PICA material property model and 4) through the evaluation of the shield material provide input to future missions which employ similar materials. Industrial X-Ray Computed Tomography (CT) is a 3D inspection technology which can provide infonnation on material integrity, material properties (density) and dimensional measurements of the heat shield components. Computed tomographic volumetric inspections can generate a dimensionally correct, quantitatively accurate volume of the shield assembly. Because of the capabilities offered by X-ray CT, NASA chose to use this method to evaluate the Stardust heat shield. Personnel at NASA Johnson Space Center (JSC) and Lawrence Livermore National Labs (LLNL) recently performed a full scan of the Stardust heat shield using a newly installed X-ray CT system at JSC. This paper briefly discusses the technology used and then presents the following results: 1. CT scans derived dimensions and their comparisons with as-built dimensions anchored with data obtained from samples cut from the heat shield; 2. Measured density variation, char layer thickness, recession and bond line (the adhesive layer between the PICA and the aeroshell) integrity; 3. FIAT predicted recession, density and char layer profiles as well as bondline temperatures Finally suggestions are made as to future uses of this technology as a tool for non-destructively inspecting and verifying both pre and post flight heat shields.

Description: The Orion Crew Module is a component of NASAs Multi-Purpose Crew Vehicle that will be used for future missions to low Earth orbit and beyond. Ten water impact tests of the Orion Ground Test Article (GTA) were conducted at the Hydro Impact Basin at NASA Langley Research Center in 2016 and were designed to provide data for the validation of the LS-DYNA model used to determine the Crew Module structural loads during ocean splashdown, and the determination of an acceptable Model Uncertainty Factor to apply to simulation results used to drive the design. Post-test data obtained from the onboard sensors were used to reconstruct the GTA trajectories both before and after water impact. Results from one vertical test and two swing tests are presented and compared to videos taken for each test.

Description: Freezable radiators offer an attractive solution to the issue of thermal control system scalability. As thermal environments change, a freezable radiator will effectively scale the total heat rejection it is capable of as a function of the thermal environment and flow rate through the radiator. Scalable thermal control systems are a critical technology for spacecraft that will endure missions with widely varying thermal requirements. These changing requirements are a result of the space craft s surroundings and because of different thermal loads during different mission phases. However, freezing and thawing (recovering) a radiator is a process that has historically proven very difficult to predict through modeling, resulting in highly inaccurate predictions of recovery time. This paper summarizes efforts made to correlate a Thermal Desktop (TM) model with empirical testing data from two test articles. A 50-50 mixture of DowFrost HD and water is used as the working fluid. Efforts to scale this model to a full scale design, as well as efforts to characterize various thermal control fluids at low temperatures are also discussed.

Description: The Internal Active Thermal Control System (IATCS) aboard the International Space Station (ISS) is primarily responsible for the removal of heat loads from payload and system racks. The IATCS is a water based system which works in conjunction with the EATCS (External ATCS), an ammonia based system, which are interfaced through a heat exchanger to facilitate heat transfer. On-orbit issues associated with the aqueous coolant chemistry began to occur with unexpected increases in CO2 levels in the cabin. This caused an increase in total inorganic carbon (TIC), a reduction in coolant pH, increased corrosion, and precipitation of nickel phosphate. These chemical changes were also accompanied by the growth of heterotrophic bacteria that increased risk to the system and could potentially impact crew health and safety. Studies were conducted to select a biocide to control microbial growth in the system based on requirements for disinfection at low chemical concentration (effectiveness), solubility and stability, material compatibility, low toxicity to humans, compatibility with vehicle environmental control and life support systems (ECLSS), ease of application, rapid on-orbit measurement, and removal capability. Based on these requirements, ortho-phthalaldehyde (OPA), an aromatic dialdehyde compound, was selected for qualification testing. This paper presents the OPA qualification test results, development of hardware and methodology to safely apply OPA to the system, development of a means to remove OPA, development of a rapid colorimetric test for measurement of OPA, and the OPA on-orbit performance for controlling the growth of microorganisms in the ISS IATCS since November 3, 2007.

Description: In the design and development of complex spacecraft missions, project teams frequently assume the use of advanced technology systems or heritage systems to enable a mission or reduce the overall mission risk and cost. As projects proceed through the development life cycle, increasingly detailed knowledge of the advanced and heritage systems within the spacecraft and mission environment identifies unanticipated technical issues. Resolving these issues often results in cost overruns and schedule impacts. The National Aeronautics and Space Administration (NASA) Discovery & New Frontiers (D&NF) Program Office at Marshall Space Flight Center (MSFC) recently studied cost overruns and schedule delays for 5 missions. The goal was to identify the underlying causes for the overruns and delays, and to develop practical mitigations to assist the D&NF projects in identifying potential risks and controlling the associated impacts to proposed mission costs and schedules. The study found that optimistic hardware/software inheritance and technology readiness assumptions caused cost and schedule growth for all five missions studied. The cost and schedule growth was not found to be the result of technical hurdles requiring significant technology development. The projects institutional inheritance and technology readiness processes appear to adequately assess technology viability and prevent technical issues from impacting the final mission success. However, the processes do not appear to identify critical issues early enough in the design cycle to ensure project schedules and estimated costs address the inherent risks. In general, the overruns were traceable to: an inadequate understanding of the heritage system s behavior within the proposed spacecraft design and mission environment; an insufficient level of development experience with the heritage system; or an inadequate scoping of the systemwide impacts necessary to implement an advanced technology for space flight applications. The paper summarizes the study s lessons learned in more detail and offers suggestions for improving the project s ability to identify and manage the technology and heritage risks inherent in the design solution.

Description: NASA s Constellation Program (CxP) was developed to successfully return humans to the Lunar surface prior to 2020. The CxP included several different project offices including Altair, which was planned to be the next generation Lunar Lander. The Altair missions were architected to be quite different than the Lunar missions accomplished during the Apollo era. These differences resulted in a significantly dissimilar Thermal Control System (TCS) design. The current paper will summarize the Altair mission architecture and the various operational phases associated with the planned mission. In addition, the derived thermal requirements and the TCS designed to meet these unique and challenging thermal requirements will be presented. During the past year, the design team has focused on developing a vehicle architecture capable of accessing the entire Lunar surface. Due to the widely varying Lunar thermal environment, this global access requirement resulted in major changes to the thermal control system architecture. These changes, and the rationale behind the changes, will be detailed throughout the current paper.

Description: Improving structural efficiency while reducing manufacturing costs are key objectives when making future heavy-lift launchers more performing and cost efficient. The main enabling technologies are the application of advanced high performance materials as well as cost effective manufacture processes. This paper presents the status and main results of a joint industrial research & development effort to demonstrate TRL 6 of a novel manufacturing process for large liquid propellant tanks for launcher applications. Using high strength aluminium-lithium alloy combined with the spin forming manufacturing technique, this development aims at thinner wall thickness and weight savings up to 25% as well as a significant reduction in manufacturing effort. In this program, the concave spin forming process is used to manufacture tank domes from a single flat plate. Applied to aluminium alloy, this process allows reaching the highest possible material strength status T8, eliminating numerous welding steps which are typically necessary to assemble tank domes from 3D-curved panels. To minimize raw material costs for large diameter tank domes for launchers, the dome blank has been composed from standard plates welded together prior to spin forming by friction stir welding. After welding, the dome blank is contoured in order to meet the required wall thickness distribution. For achieving a material state of T8, also in the welding seams, the applied spin forming process allows the required cold stretching of the 3D-curved dome, with a subsequent ageing in a furnace. This combined manufacturing process has been demonstrated up to TRL 6 for tank domes with a 5.4 m diameter. In this paper, the manufacturing process as well as test results are presented. Plans are shown how this process could be applied to future heavy-lift launch vehicles developments, also for larger dome diameters.

Description: Node 1 flew to the International Space Station (ISS) on Flight 2A during December 1998. To date the National Aeronautics and Space Administration (NASA) has learned a lot of lessons from this module based on its history of approximately two years of acceptance testing on the ground and currently its twelve years on-orbit. This paper will provide an overview of the ISS Environmental Control and Life Support (ECLS) design of the Node 1 Temperature and Humidity Control (THC) subsystem and it will document some of the lessons that have been learned to date for this subsystem and it will document some of the lessons that have been learned to date for these subsystems based on problems prelaunch, problems encountered on-orbit, and operational problems/concerns. It is hoped that documenting these lessons learned from ISS will help in preventing them in future Programs. 1

Description: The Hayabusa (originally known as MUSES-C) engineering spacecraft was launched by the 5th Mu V launch vehicle on May 9, 2003 by the Japan Aerospace Exploration Agency (JAXA). It was designed to acquire samples from the surface of near-Earth asteroid 25143 Itokawa (1998 SF36) and return them to Earth. The main objectives of the mission were to demonstrate the performance of various technologies such as ion engine performance, autonomous navigation and control, asteroid surface sampling, and recovery of the return capsule after high speed re-entry. Hayabusa successfully returned a small capsule to Earth in June 2010 with a parachute assisted landing in Woomera, Australia. Details of the Hayabusa mission and the recovery operation will be presented for discussion.

Description: Many earth observing sensors depend on white diffuse reflectance standards to derive scales of radiance traceable to the St Despite the large number of Earth observing sensors that operate in the reflective solar region of the spectrum, there has been no direct method to provide NIST traceable BRDF measurements out to 2500 rim. Recent developments in detector technology have allowed the NIST reflectance measurement facility to expand the operating range to cover the 250 nm to 2500 nm range. The facility has been modified with and additional detector using a cooled extended range indium gallium arsenide (Extended InGaAs) detector. Measurements were made for two PTFE white diffuse reflectance standards over the 1100 nm to 2500 nm region at a 0' incident and 45' observation angle. These two panels will be used to support the OLI calibration activities. An independent means of verification was established using a NIST radiance transfer facility based on spectral irradiance, radiance standards and a diffuse reflectance plaque. An analysis on the results and associated uncertainties will be discussed.

Description: No abstract available

Description: The International Space Station (ISS) has established a new model for the achievement of the most difficult engineering goals in space: international collaboration at the program level with competition at the level of technology. This strategic shift in management approach provides long term program stability while still allowing for the flexible evolution of technology needs and capabilities. Both commercial and government sponsored technology developments are well supported in this management model. ISS also provides a physical platform for development and demonstration of the systems needed for missions beyond low earth orbit. These new systems at the leading edge of technology require operational exercise in the unforgiving environment of space before they can be trusted for long duration missions. Systems and resources needed for expeditions can be aggregated and thoroughly tested at ISS before departure thus providing wide operational flexibility and the best assurance of mission success. We will describe representative mission profiles showing how ISS can support exploration missions to the Moon, Mars, asteroids and other potential destinations. Example missions would include humans to lunar surface and return, and humans to Mars orbit as well as Mars surface and return. ISS benefits include: international access from all major launch sites; an assembly location with crew and tools that could help prepare departing expeditions that involve more than one launch; a parking place for reusable vehicles; and the potential to add a propellant depot.

Description: Vulnerability of a variety of candidate spacecraft electronics to total ionizing dose and displacement damage is studied. Devices tested include optoelectronics, digital, analog, linear bipolar devices, and hybrid devices.

Description: The Soil Moisture Active Passive (SMAP) mission is a NASA directed mission to map global land surface soil moisture and freeze-thaw state. Instrument and mission details are shown. The key SMAP soil moisture product is provided at 10 km resolution with 0.04cubic cm/cubic cm accuracy. The freeze/thaw product is provided at 3 km resolution and 80% frozen-thawed classification accuracy. The full list of SMAP data products is shown.

Description: A Technology Computer Aided Design (TCAD) simulation-based method is developed to evaluate whether derating of high-energy heavy-ion accelerator test data bounds the risk for single-event gate rupture (SEGR) from much higher energy on-orbit ions for a mission linear energy transfer (LET) requirement. It is shown that a typical derating factor of 0.75 applied to a single-event effect (SEE) response curve defined by high-energy accelerator SEGR test data provides reasonable on-orbit hardness assurance, although in a high-voltage power MOSFET, it did not bound the risk of failure.

Description: The NASA Lunar Electric Rover (LER) has been developed at Johnson Space Center as a next generation mobility platform. Based upon a twelve wheel omni-directional chassis with active suspension the LER introduces a number of novel capabilities for lunar exploration in both manned and unmanned scenarios. Besides being the primary vehicle for astronauts on the lunar surface, LER will perform tasks such as lunar regolith handling (to include dozing, grading, and excavation), equipment transport, and science operations. In an effort to support these additional tasks a team at the Kennedy Space Center has produced a universal attachment interface for LER known as the Quick Attach. The Quick Attach is a compact system that has been retro-fitted to the rear of the LER giving it the ability to dock and undock on the fly with various implements. The Quick Attach utilizes a two stage docking approach; the first is a mechanical mate which aligns and latches a passive set of hooks on an implement with an actuated cam surface on LER. The mechanical stage is tolerant to misalignment between the implement and the LER during docking and once the implement is captured a preload is applied to ensure a positive lock. The second stage is an umbilical connection which consists of a dust resistant enclosure housing a compliant mechanism that is optionally actuated to mate electrical and fluid connections for suitable implements. The Quick Attach system was designed with the largest foreseen input loads considered including excavation operations and large mass utility attachments. The Quick Attach system was demonstrated at the Desert Research And Technology Studies (D-RA TS) field test in Flagstaff, AZ along with the lightweight dozer blade LANCE. The LANCE blade is the first implement to utilize the Quick Attach interface and demonstrated the tolerance, speed, and strength of the system in a lunar analog environment.

Description: Free-piston Stirling convertors are fundamental to the development of NASA s next generation of radioisotope power system, the Advanced Stirling Radioisotope Generator (ASRG). The ASRG will use General Purpose Heat Source (GPHS) modules as the energy source and Advanced Stirling Convertors (ASCs) to convert heat into electrical energy, and is being developed by Lockheed Martin under contract to the Department of Energy. Achieving flight status mandates that the ASCs satisfy design as well as flight requirements to ensure reliable operation during launch. To meet these launch requirements, GRC performed a series of quasi-static mechanical tests simulating the pressure, thermal, and external loading conditions that will be experienced by an ASC-E2 heater head assembly. These mechanical tests were collectively referred to as "lateral load tests" since a primary external load lateral to the heater head longitudinal axis was applied in combination with the other loading conditions. The heater head was subjected to the operational pressure, axial mounting force, thermal conditions, and axial and lateral launch vehicle acceleration loadings. To permit reliable prediction of the heater head s structural performance, GRC completed Finite Element Analysis (FEA) computer modeling for the stress, strain, and deformation that will result during launch. The heater head lateral load test directly supported evaluation of the analysis and validation of the design to meet launch requirements. This paper provides an overview of each element within the test and presents assessment of the modeling as well as experimental results of this task.

Description: Dust accumulation on thermal radiator surfaces planned for lunar exploration will significantly reduce their efficiency. Evidence from the Apollo missions shows that an insulating layer of dust accumulated on radiator surfaces could not be removed and caused serious thermal control problems. Temperatures measured at different locations in the magnetometer on Apollo 12 were 38 C warmer than expected due to lunar dust accumulation. In this paper, we report on the application of the Electrodynamic Dust Shield (EDS) technology being developed in our NASA laboratory and applied to thermal radiator surfaces. The EDS uses electrostatic and dielectrophoretic forces generated by a grid of electrodes running a 2 micro A electric current to remove dust particles from surfaces. Working prototypes of EDS systems on solar panels and on thermal radiators have been successfully developed and tested at vacuum with clearing efficiencies above 92%. For this work EDS prototypes on flexible and rigid thermal radiators were developed and tested at vacuum.

Description: Safety of the next-generation space flight vehicles requires development of an in-flight Failure Detection and Prognostic (FD&P) system. Development of such system is challenging task that involves analysis of many hard hitting engineering problems across the board. In this paper we report progress in the development of FD&P for the re-contact fault between upper stage nozzle and the inter-stage caused by the first stage and upper stage separation failure. A high-fidelity models and analytical estimations are applied to analyze the following sequence of events: (i) structural dynamics of the nozzle extension during the impact; (ii) structural stability of the deformed nozzle in the presence of the pressure and temperature loads induced by the hot gas flow during engine start up; and (iii) the fault induced thrust changes in the steady burning regime. The diagnostic is based on the measurements of the impact torque. The prognostic is based on the analysis of the correlation between the actuator signal and fault-induced changes in the nozzle structural stability and thrust.

Description: A series of tests were conducted to evaluate protuberance heating for the purposes of vehicle design and modification. These tests represent a state of the art approach to both testing and instrumentation for defining aerothermal protuberance effects on the protuberance and surrounding areas. The testing was performed with a number of wind tunnel entries beginning with the proof of concept "pathfinder" test in the Test Section 1 (TS1) tunnel in the Langley Unitary Plan Wind Tunnel (UPWT). The TS1 section (see Figures 1a and 1b) is a lower Mach number tunnel and the Test Section 2 (TS2) has overlapping and higher Mach number capability as showin in Figure 1c. The pathfinder concept was proven and testing proceeded for a series of protuberance tests using an existing splitter aluminum protuberance mounting plate, Macor protuberances, thin film gages, total temperature and pressure gages, Kulite pressure transducers, Infra-Red camera imaging, LASER velocimetry evaluations and the UPWT data collection system. A boundary layer rake was used to identify the boundary layer profile at the protuberance locations for testing and helped protuberance design. This paper discusses the techniques and instrumentation used during the protuberance heating tests performed in the UPWT in TS1 and TS2. Runs of the protuberances were made Mach numbers of 1.5, 2.16, 2.65, and 3.51. The data set generated from this testing is for ascent protuberance effects and is termed Protuberance Heating Ascent Data (PHAD) and this testing may be termed PHAD-1 to distinguish it from future testing of this type.

Description: Polymers and other oxidizable materials on the exterior of spacecraft in the low Earth orbit (LEO) space environment can be eroded due to reaction with atomic oxygen (AO). Therefore, in order to design durable spacecraft, it is important to know the LEO AO erosion yield (Ey, volume loss per incident oxygen atom) of materials susceptible to AO reaction. The Polymers Experiment was developed to determine the AO Ey of various polymers and other materials flown in ram and wake orientations in LEO. The experiment was flown as part of the Materials International Space Station Experiment 7 (MISSE 7) mission for 1.5 years on the exterior of the International Space Station (ISS). As part of the experiment, a sample containing Class 2A diamond (100 plane) and highly oriented pyrolytic graphite (HOPG, basal and edge planes) was exposed to ram AO and characterized for erosion. The materials were salt-sprayed prior to flight to provide isolated sites of AO protection. The Ey of the samples was determined through post-flight electron microscopy recession depth measurements. The experiment also included a Kapton H witness sample for AO fluence determination. This paper provides an overview of the MISSE 7 mission, a description of the flight experiment, the characterization techniques used, the mission AO fluence, and the LEO Ey results for diamond and HOPG (basal and edge planes). The data is compared to the Ey of pyrolytic graphite exposed to four years of space exposure as part of the MISSE 2 mission. The results indicate that diamond erodes, but with a very low Ey of 1.58 +/- 0.04 x 10(exp -26) cm(exp 3)/atom. The different HOPG planes displayed significantly different amounts of erosion from each other. The HOPG basal plane had an Ey of 1.05 +/- 0.08 x 10(exp -24) cm(exp 3)/atom while the edge plane had a lower Ey of only 5.38 +/- 0.90 x 10(exp -25) -cm(exp 3)/atom. The Ey data from this ISS spaceflight experiment provides valuable information for understanding of chemistry and chemical structure dependent modeling of AO erosion.

Description: Current concepts of operations for human exploration of Mars center on the staged deployment of spacecraft, logistics, and crew. Though most studies focus on the needs for human occupation of the spacecraft and habitats, these resources will spend most of their lifetime unoccupied. As such, it is important to identify the operational state of the unoccupied spacecraft or habitat, as well as to design the systems to enable the appropriate level of autonomy. Key goals for this study include providing a realistic assessment of what "dormancy" entails for human spacecraft, exploring gaps in state-of-the-art for autonomy in human spacecraft design, providing recommendations for investments in autonomous systems technology development, and developing architectural requirements for spacecraft that must be autonomous during dormant operations. The mission that was chosen is based on a crewed mission to Mars. In particular, this study focuses on the time that the spacecraft that carried humans to Mars spends dormant in Martian orbit while the crew carries out a surface mission. Communications constraints are assumed to be severe, with limited bandwidth and limited ability to send commands and receive telemetry. The assumptions made as part of this mission have close parallels with mission scenarios envisioned for dormant cis-lunar habitats that are stepping-stones to Mars missions. As such, the data in this report is expected to be broadly applicable to all dormant deep space human spacecraft.

Description: As spacecraft travel through space plasma, spacecraft surfaces become charged by the collection of charged particles. This process is referred to as Surface Charging. These charges can be detrimental to the vehicle's electronic subsystems as they present a threat of electrostatic discharge (ESD) to onboard circuitry. The process of Surface Charging is complex and is affected by many elements. The charging of each surface is unique. The potential of an individual surface is dependent upon many variables including but not limited to the surface's geometry, material and its location. Each surface also has unique interactions with the surrounding plasma. Other factors that play large roles in the charging process is the density and temperature of plasma ions and electrons. Using Nascap-2k, a model of the Freja satellite has been constructed, and its auroral plasma environment has been imitated to simulate surface charging characteristics. The charging process of the Freja satellite has been modeled iteratively with incremental changes in both the Maxwellian electron temperature (eV) as well as the Gaussian electron energy (eV). This study provides an analysis of the sensitivity between spacecraft surface charging and these two primary variables of electron differential flux.

Description: As spacecraft travel through plasma, spacecraft surfaces become charged by the collection of charged particles. This process is referred to as Surface Charging. These charges can be detrimental to the vehicle's electronic subsystems as they present a threat of electrostatic discharge (ESD) to onboard circuitry. The process of Surface Charging is complex and is affected by many elements. The charging of each surface is unique. The potential of an individual surface is dependent upon many variables including but not limited to the surface's geometry, material and its location. Each surface also has unique interactions with the surrounding plasma. Other factors that play large roles in the charging process is the density and temperature of plasma ions and electrons. Using Nascap-2k, a model of the Freja satellite has been constructed, and its auroral plasma environment has been imitated to simulate surface charging characteristics. The charging process of the Freja satellite has been modeled iteratively with incremental changes in both the Maxwellian electron temperature (eV) as well as the Gaussian electron energy (eV). This study provides an analysis of the sensitivity between spacecraft surface charging and these two primary variables of electron differential flux.

Description: CubeSats are a relatively new type of satellite. Smaller than long-term (5+ year life expectancy) satellites, these pico-satellites are comparatively cheap, small (10x10x10 cm), and are very versatile. Universities world-wide are using CubeSats to conduct a variety of experiments in space without the need for a large experimental platform. Today CubeSats are considered to be one of the most effective ways to send a small payload into space and has attracted the attention of many educational and non-profit organizations. As this pico-satellite model continues to gain penetration into the satellite build and launch industry, it is expected that more governmental, educational, and commercial interests will emerge. As an example, more of the space-related items of high interest to the National Science Foundation may be tackled with a CubeSat platform resulting in lower life cycle costs than traditional satellite options. NASA LSP, in cooperation with the Florida Institute of Technology, has initiated a feasibility study to investigate the technical aspects of measuring and transferring vibration, acceleration, temperature, and video data from a CubeSat to NASA Hanger AE on Cape Canaveral Air Force Station (CCAFS) a.k.a. Kennedy Space Center (KSC). This report provides a technical feasibility analysis to determine whether-or-not a specific set of NASA/LSP requirements can be accomplished. Our approach has been to provide a "notional" component layout to determine the feasibility of the NASA/LSP stakeholder requirements. The notional layout is used to consider component level technical issues such as size, weight, & power (SWaP), bandwidth, and other critical technical parameters. Even though the notional components may satisfy the stated requirements and thereby demonstrate feasibility, the notional layout is NOT considered a design since no component optimization and design trade-off analysis has taken place. This activity should be accomplished in an appropriate design phase that is outside of the scope of this effort.

Description: Alkali liquid metal cooled fission reactor concepts are under development for mid-range spaceflight power requirements. One such concept utilizes a sodium-potassium eutectic (NaK) as the primary loop working fluid. Traditionally, linear induction pumps have been used to provide the required flow and head conditions for liquid metal systems but can be limited in performance. This paper details the design, build, and check-out test of a mechanical NaK pump. The pump was designed to meet reactor cooling requirements using commercially available components modified for high temperature NaK service.

Description: The Flight Loads Laboratory at the Dryden Flight Research Center conducted tests to measure the inertia properties of the Orion Pad Abort 1 (PA-1) Crew Module. These measurements were taken to validate analytical predictions of the inertia properties of the vehicle and assist in reducing uncertainty for derived aero performance results calculated post launch. The first test conducted was to determine the Ixx of the Crew Module. This test approach used a modified torsion pendulum test step up that allowed the suspended Crew Module to rotate about the x axis. The second test used a different approach to measure both the Iyy and Izz properties. This test used a Knife Edge fixture that allowed small rotation of the Crew Module about the y and z axes. Discussions of the techniques and equations used to accomplish each test are presented. Comparisons with the predicted values used for the final flight calculations are made. Problem areas, with explanations and recommendations where available, are addressed. Finally, an evaluation of the value and success of these techniques to measure the moments of inertia of the Crew Module is provided.

Description: Sublimators have been used as heat rejection devices for a variety of space applications including the Apollo Lunar Module and the Extravehicular Mobility Unit (EMU). Sublimators typically operate with steady-state feedwater utilization at or near 100%. However, sublimators are currently being considered to operate in a cyclical topping mode, which represents a new mode of operation for sublimators. Sublimators can be used as a topper during mission phases such as low lunar or low earth orbit. In these mission phases, the sublimator will be repeatedly started and stopped during each orbit to provide supplemental heat rejection for the portion of the orbit where the radiative sink temperature exceeds the system setpoint temperature. This paper will investigate the effects of these transient starts and stops on the feedwater utilization during various feedwater timing scenarios. The X-38 sublimator and Contamination Insensitive Sublimator (CIS) were tested in a ground vacuum chamber to understand this behavior and to quantify the feedwater performance. Data from various scenarios will be analyzed to investigate feedwater utilization under the cyclical conditions. This paper will also provide recommendations for future sublimator designs and/or feedwater control.

Description: Following recommendations by the National Research Council, NASA's Authorization Act of 2008 (P.I. 110-422) and the Fiscal Year 2009 Omnibus Appropriations Act directed NASA to assess the feasibility of using the planned human spaceflight architecture to service existing and future observatory-class scientific spacecraft. This interest in space servicing, either with astronauts and/or with robots, reflects the decades-long success that NASA has achieved with the Space Shuttle program and the Hubble Space Telescope on behalf of the international astronomical community. This study is led by NASA Goddard Space Flight Center and will last about a year, leading to an assessment report to NASA and the science communities. We will report on the status of this study, progress toward goals, workshops, and priorities for the next few months.

Description: It has been suggested that the International Space Station (ISS) be utilized to simulate the transit portion of long-duration missions to Mars and near-Earth asteroids (NEA). The ISS offers a unique environment for such simulations, providing researchers with a high-fidelity platform to study, enhance, and validate technologies and countermeasures for these long-duration missions. From a space life sciences perspective, two major categories of human research activities have been identified that will harness the various capabilities of the ISS during the proposed simulations. The first category includes studies that require the use of the ISS, typically because of the need for prolonged weightlessness. The ISS is currently the only available platform capable of providing researchers with access to a weightless environment over an extended duration. In addition, the ISS offers high fidelity for other fundamental space environmental factors, such as isolation, distance, and accessibility. The second category includes studies that do not require use of the ISS in the strictest sense, but can exploit its use to maximize their scientific return more efficiently and productively than in ground-based simulations. In addition to conducting Mars and NEA simulations on the ISS, increasing the current increment duration on the ISS from 6 months to a longer duration will provide opportunities for enhanced and focused research relevant to long-duration Mars and NEA missions. Although it is currently believed that increasing the ISS crew increment duration to 9 or even 12 months will pose little additional risk to crewmembers, additional medical monitoring capabilities may be required beyond those currently used for the ISS operations. The use of the ISS to simulate aspects of Mars and NEA missions seems practical, and it is recommended that planning begin soon, in close consultation with all international partners.

Description: A wealth of literature exists on control allocation algorithms for over-actuated air vehicles, launch vehicles, and spacecraft's. Most of these algorithms focus primarily on minimizing some objective function such as command tracking error and/or control effector usage. Linear allocators (pseudo inverses) are usually the conventional choice due to their simplicity and the ability to achieve a significant portion of the theoretical moment/impulse space. Generally, it is assumed that there exists minimal interaction effects between control effectors. In fact, very few studies address the problem of control effector interactions in the context of control allocation, especially for small spacecraft's with a reaction control system (RCS). This paper presents a CubeSat RCS design with a four thruster tetrahedral layout such that when two or more thrusters re, the resultant impulse differs noticeably compared to the sum of the contributions from individual thruster rings. This undesirable effect is caused by the design of the propellant tank and regulator. To mitigate this issue, an innovative modified pseudo inverse (MPI) control allocation algorithm was developed that adjusts the pseudo inverse solution based on test data. The algorithm is iteration-free and superior to the standard pseudo inverse in minimizing the command tracking error.

Description: Here we describe the Primitive Object Volatile Explorer (PrOVE), a smallsat mission concept to study the surface structure and volatile inventory of comets in their perihelion passage phase when volatile activity is near peak. CubeSat infrastructure imposes limits on propulsion systems, which are compounded by sensitivity to the spacecraft disposal state from the launch platform and potential launch delays. We propose circumventing launch platform complications by using waypoints in space to park a deep space SmallSat or CubeSat while awaiting the opportunity to enter a trajectory to flyby a suitable target. In our Planetary Science Deep Space SmallSat Studies (PSDS3) project, we investigated scientific goals, waypoint options, potential concept of operations (ConOps) for periodic and new comets, spacecraft bus infrastructure requirements, launch platforms, and mission operations and phases. Our payload would include two low-risk instruments: a visible image (VisCAM) for 5-10 m resolution surface maps; and a highly versatile multispectral Comet CAMera (ComCAM) will measure 1) H2O, CO2, CO, and organics non-thermal fluorescence signatures in the 2-5 m MWIR, and 2) 7-10 and 8-14 m thermal (LWIR) emission. This payload would return unique data not obtainable from ground-based telescopes and complement data from Earth-orbiting observatories. Thus, the PrOVE mission would (1) acquire visible surface maps, (2) investigate chemical heterogeneity of a comet nucleus by quantifying volatile species abundance and changes with solar insolation, (3) map the spatial distribution of volatiles and determine any variations, and (4) determine the frequency and distribution of outbursts.

Description: No abstract available

Description: Much effort has been made to enhance exploration on Mars. In addition to a rover and Mars-orbiting satellites, a Mars helicopter (MH) was proposed in order to augment planetary research. Computational Fluid Dynamics (CFD) simulations have been performed to have a better understanding of the behavior and performance of vertical lift Planetary Aerial Vehicles (PAV). Due to the large differences in atmospheric conditions between Mars and Earth, predicting and testing rotorcraft performance is a complex task. The goal of this project is to understand the capability of the mid-fidelity CFD software RotCFD to predict rotor performance in terms of thrust at 1013.25 milibar and 14 milibar corresponding to Terrestrial and Martian conditions, respectively. Also, in order to characterize the wind tunnel wall effects free field and wind tunnel simulations were performed, analyzed and compared. Different analytical tools have been used in order to aid with the design process for the future vertical lift planetary aerial vehicles. One of them includes experimental tests performed on a rotor in the Aeolian Wind Tunnel (AWT) facility at NASA Ames Research Center under different pressure conditions ranging from Terrestrial to Martian atmospheric conditions. Other software was used as well in order to capture the aerodynamic coefficients of the corresponding rotor sections based on the Mach and Reynolds numbers used for the experimental tests. The aerodynamic coefficients were input into RotCFD, and various simulations were performed under Terrestrial and Martian conditions in order to mimic the experimental test. Then, the obtained results from RotCFD were compared with the AWT collected data.

Description: Aerocapture has been extensively studied and these studies have shown the benefit for planetary exploration missions. While the traditional approach to aerocapture with lifting configurations and lift-guided modulations have been assessed to be technologically feasible, aerocapture using purely drag modulation was proposed and studied by Prof. Braun and his students. These studies show that if one can assess the feasibility of aerocapture using drag modulation at Venus, and develop tall pole technologies needed at Venus, then this concept is much easier to execute at all other relevant destinations. Based on the above finding, partnered proposals were submitted by Adam Nelessen at JPL and Ethiraj Venkatapathy at Ames in collaboration with Prof. Braun at the University of Colorado, Boulder (UCB). Under this partnership, Ames Research Center (ARC) is working to address some of the key entry technology challenges associated with drag modulation aerocapture at Venus. Drag modulation aerocapture is a simple, scalable, and likely cost-effective way to enhance planetary science missions. The approach envisioned is to design a small spacecraft, that would most likely be a secondary payload, with a removable drag skirt. The vehicle would enter the atmosphere at Venus with a low ballistic coefficient, decelerate rapidly, drop the skirt resulting in a smaller vehicle with a higher ballistic coefficient which would skip out of the atmosphere and enter into a desired orbit. ARC's role in this collaboration is multifold. First of which is to perform design studies on various pre- and post-jettison geometries utilizing a 3-DOF trajectory code to determine the aerodynamics and aerothermodynamics of the vehicles and evaluate viable thermal protection material system designs. Once these design studies are complete, Ames will then perform higher fidelity CFD and TPS sizing to further design the vehicles. Second, the multi-body separation dynamics of the drag modulation event will be explored using both CFD simulations (CART3-D and US3D) as well as possible ballistic range testing. ARC's tools and expertise have been used to assess and advise on the selection of the separating configuration. In addition to the preliminary evaluation, ARC will provide tools and expertise to UCB team members to further assess aerodynamic interactions between the separating bodies and provide guidance as to the feasibility of stable transition.

Description: No abstract available

Description: The NASA Glenn Research Center (GRC) in Cleveland, Ohio designs and develops innovative technologies to advance NASA's missions in aeronautics and space exploration. The center's expertise includes that in power, energy storage, and conversion; in-space chemical and electric propulsion; communications; and instrumentation technologies. GRC is currently managing and/or developing a number of these technologies for Small Spacecraft applications. Small spacecraft propulsion efforts include efforts with Tethers Unlimited, Inc. (TUI) and Busek. Power systems technology efforts include the Advanced Electrical Bus (ALBus) CubeSat inhouse development as well as efforts with Rochester Institute of Technology (RIT), the Kennedy Space Center & the University Miami. In the area of communications, NASA-GRC continues to explore the potential capabilities and advantages of using Ka-band for LEO (Low Earth Orbit) spacecraft communications with both NASA and commercially owned GEO (Geosynchrous Earth Orbit) relays and direct-to-ground terminal networks. GRC has also proposed a number of small spacecraft instrumentation technology demonstration such as SPAGHETI (Solar Proton Anisotropy and Galactic cosmic ray High Energy Transport Instrument) and CFIDS (Compact Full-Field Ion Detector System).

Description: No abstract available

Description: A regeneratively-cooled nozzle for liquid rocket engine applications is a significant cost of the overall engine due to the complexities of manufacturing a large thin-walled structure that must operate in extreme temperature and pressure environments. NASA has been investigating and advancing methods for fabrication of liquid rocket engine channel wall nozzles to realize further cost and schedule improvements. The methods being evaluated are targeting increased scale required for current NASA and commercial space programs. Several advanced rapid fabrication methods are being investigated for forming of the inner liner, producing the coolant channels, closeout of the coolant channels, and fabrication of the manifolds. NASA Marshall Space Flight Center (MSFC) completed process development and subscale hot-fire testing of a series of these advanced fabrication channel wall nozzle technologies to gather performance data in a relevant environment. The primary fabrication technique being discussed in this paper is Laser Wire Deposition Closeout (LWDC). This process has been developed to significantly reduce time required for closeouts of regeneratively-cooled slotted liners. It allows for channel closeout to be formed in place in addition to the structural jacket without the need for channel fillers or complex tooling. Additional technologies were also tested as part of this program including water jet milling and arc-based additive manufacturing deposition. Each nozzle included different fabrication features, materials, and methods to demonstrate durability in a hot-fire environment. The results of design, fabrication and hot-fire testing performance is discussed in this paper.

Description: The Space Launch System (SLS) Block-1B vehicle includes a low thrust-to-weight upper stage, which presents challenges to heritage ascent guidance algorithms. A trade study was conducted to evaluate two alternative guidance algorithms: 1) Powered Explicit Guidance (PEG), based on a modified implementation of PEG used on the Block-1 vehicle, and 2) Optimal Guidance (OPGUID), an algorithm developed for Marshall Space Flight Center (MSFC) and used on Constellation and other Guidance, Navigation, and Controls (GN&C) projects. The design criteria, approach, and results of the trade study are given, as well as other impacts and considerations for Block-1B type missions.

Description: This paper details the results of an initial study to develop a certification plan for human-rated inflatable space structures, including guidelines for qualification testing. Habitable softgoods inflatables are multi-layered shell structures that use high-strength webbing, cordage and broadcloth fabric to carry the skin loads of a variety of volumetric shapes and structural architectures. The primary objectives of this study are to define the key parameters that affect these structures and propose a statistically robust approach to defining safety and knockdown factors based on test and analysis. Current NASA standards for habitable inflatable space structures use a factor of safety of 4, which was inherited from airship design criteria. An updated approach to defining a design factor, taking into account material strength variability, load variability in the article, number of test samples, and damage and degradation effects is specified. Accurate analytical modeling of these structures is hindered by the difficulty of obtaining accurate and consistent material data due to load-history- dependent, nonlinear load versus strain behavior. A building block approach to certification is detailed that uses stochastic modeling and statistical test design and analysis to address the unique challenges these high-strength softgoods structures present. Human-rated inflatable modules are a transformative capability for launching much larger habitable volumes into space than is possible with rigid shell structures. This research aims to provide the framework for certifying these structures for future human space exploration missions.

Description: Final Document is attached. The Robotic External Leak Locator (RELL) was deployed to the International Space Station (ISS) with the goal of detecting and locating on-orbit leaks around the ISS. Three activities to investigate and corroborate the background natural and induced environment of ISS were performed with RELL as part of the on-orbit validation and demonstration conducted in November December 2016. The first demonstration activity pointed RELL directly in the ram and wake directions for one orbit each. The ram facing measurements showed high partial pressure for mass-to-charge ratio 16, corresponding to atomic oxygen (AO), as well as the presence of mass-to-charge ratio 17. RELLs view in the wake-facing direction included more ISS structure and several Environmental Control and Life Support System (ECLSS) on-orbit vents were detected, including the Carbon Dioxide Removal Assembly (CDRA), Russian segment ECLSS, and Sabatier vents. The second demonstration activity pointed RELL at three faces of the P1 Truss segment. Effluents from ECLSS and European Space Agency (ESA) Columbus module on-orbit vents were detected by RELL. The partial pressures of mass-to-charge ratios 17 and 18 remained consistent with the first on-orbit activity of characterizing the natural environment. The third demonstration activity involved RELL scanning an Active Thermal Control System (ATCS) radiator. Three locations along the radiator were scanned and the angular position of RELL with respect to the radiator was varied. Mass-to-charge ratios 16 and 17 both had upward shifts in partial pressure when pointing toward the Radiator Beam Valve Modules (RBVMs), likely corresponding to a known, small ammonia leak.

Description: The 3rd Planetary CubeSat Science Symposium will be held at NASA Goddard Space Flight Center, with the participation of CubeSat/SmallSat scientists and developers. Discussions will include current missions, mission concepts, and opportunities for future mission selections. The sessions will also include panel discussions about strategic and technical aspects of planetary small satellite missions, and an afternoon poster session providing mission proposers the opportunity to meet with vendors and suppliers. This presentation (no paper), will provide an overview of the navigation systems avaiable for Cubesat Planetary missions.

Description: Phenolic Impregnated Carbon Ablator (PICA), invented in the mid 1990's, is a low-density ablative thermal protection material proven capable of meeting sample return mission needs from the moon, asteroids, comets and other "unrestricted class V destinations" as well as for Mars. Its low density and efficient performance characteristics have proven effective for use from Discovery to Flagship class missions. It is important that NASA maintain this TPS material capability and ensure its availability for future NASA use. The rayon based carbon precursor raw material used in PICA preform manufacturing required replacement and requalification at least twice in the past 25 years and a third substitution is now needed. The carbon precursor replacement challenge is twofold the first involves finding a long-term replacement for the current rayon and the second is to assess its future availability periodically to ensure it is sustainable and be alerted if additional replacement efforts need to be initiated. Rayon is no longer a viable process in the US and Europe due to environmental concerns. In the early 80's rayon producers began investigating a new method of producing a cellulosic fiber through a more environmentally responsible process. This cellulosic fiber, lyocell, is a viable replacement precursor for PICA fiberform. This presentation reviews current SMD-PSD funded PICA sustainability activities in ensuring a rayon replacement for the long term is identified and in establishing that the capability of the new PICA derived from an alternative precursor is in family with previous versions of the so called "heritage" PICA.State of the Art Low Density Carbon Phenolic AblatorsStardust SRC post flight withPICA forebody heat shield(0.8m max. diameter)PICA Processing StepsRole of Rayon/Lyocellin PICA.

Description: No abstract available

Description: No abstract available

Description: This presentation introduces a new sizing and margin methodology for dual-layer Thermal Protection Systems (TPS). The methodology has been tailored for application to a dual-layer 3D-woven TPS called Heat-shield for Extreme Entry Environments Technology (HEEET). Sizing is performed for a reference Saturn probe mission to show how uncertainties in trajectory, aerothermal modelling and TPS response impact the sizing of each layer.

Description: The Origins, Spectral Interpretation, Resource Identification, Security, Regolith Explorer (OSIRIS-REx) Visible and Infrared Spectrometer (OVIRS) is a cryogenic instrument. At the Outbound Cruise nominal spacecraft attitude, sunlight impinges on several multilayer insulation blankets on the forward deck. It is reflected or scattered to other components on the deck. This solar illumination adds heat load to the OVIRS, and causes its detector temperature to exceed the 105K maximum operating allowable flight temperature limit by 0.8K. During the flight system thermal vacuum test, the solar simulator beam reflected or scattered from the test fixtures to the OVIRS added non-flight heat load. The detector temperature was 9K warmer than that in flight. At those temperatures, the science data was acceptable, despite its quality was not as high as that of 105K or colder.

Description: This paper summarizes the on-orbit structural dynamic data and the related modal analysis, model validation and correlation performed for the International Space Station (ISS) configuration ISS Stage ULF7, 2015 Dedicated Thruster Firing (DTF). The objective of this analysis is to validate and correlate the analytical models used to calculate the ISS internal dynamic loads and compare the 2015 DTF with previous tests. During the ISS configurations under consideration, on-orbit dynamic measurements were collected using the three main ISS instrumentation systems; Internal Wireless Instrumentation System (IWIS), External Wireless Instrumentation System (EWIS) and the Structural Dynamic Measurement System (SDMS). The measurements were recorded during several nominal on-orbit DTF tests on August 18, 2015. Experimental modal analyses were performed on the measured data to extract modal parameters including frequency, damping, and mode shape information. Correlation and comparisons between test and analytical frequencies and mode shapes were performed to assess the accuracy of the analytical models for the configurations under consideration. These mode shapes were also compared to earlier tests. Based on the frequency comparisons, the accuracy of the mathematical models is assessed and model refinement recommendations are given. In particular, results of the first fundamental mode will be discussed, nonlinear results will be shown, and accelerometer placement will be assessed.

Description: The NASA Technical Fellows periodically conduct State-of-the-Discipline assessments. The GN&C Technical Fellow contracted Harlan Brown & Company in 2007 and 2009 to conduct independent, third party studies to gain unbiased insight and understanding into the attitudes and beliefs of NASA's GN&C Community of Practice (CoP). The paper first outlines the background, objectives and methodology of the studies. The paper then summarizes key study results of the 2007 baseline study, as well as the 2009 update. The update was then used to track and monitor perceptions, identify performance trends, identify areas where further improvement needs to be made in NASA's GN&C discipline. It also generated feedback on the recently developed GN&C CoP online knowledge capture and learning site.

Description: No abstract available

Description: The Mars2020 entry vehicle is currently being developed by NASA to safely land its next rover on the Martian surface in 2021. During entry, the vehicle will be protected from aerothermal environments using a PICA (Phenolic Impregnated Carbon Ablator)-tiled heatshield. PICA loses mass through surface recession and in-depth pyrolysis as it is heated. Pre-flight knowledge of heatshield mass loss is required for vehicle balancing during critical mission events. This study attempts to predict the total mass loss experienced by the Mars2020's heatshield during its entry. A grid was created over the half of the heatshield which generated 108 points across a total of 9 spokes. Aero-thermal environments were provided from CFD (Computational Fluid Dynamics) calculations that considered a baselined trajectory. The TPS (Thermal Protection System) stack was a build-up of composite, aluminum, composite, an HT-424 bond, followed by PICA. The FIAT (Fully Implicit Ablation, Thermal-response) 1-D analysis utilized this TPS stack and the CFD environments and was run at each grid point giving mass flux information from the point of atmospheric entry until parachute deployment. The mass flux due to recession and pyrolysis gas was summed and integrated first through time and then across the half heatshield using a polar integration tool. The mass loss results were mirrored to the other half of the heatshield to calculate total mass loss throughout the entry phase of flight. This total mass loss value and its distribution was used by entry vehicle designers to account for CG (Center of Gravity) offset during parachute descent when the heatshield is no longer losing significant mass.

Description: This paper presents an overview of the design optimisation measures that have been proposed and analysed in order to reduce the mass of the structure, including the MMOD (Micro-Meteoroid and Orbital Debris) protection system, of the ESM (European Service Module) for the Orion MPCV (Multi-Purpose Crew Vehicle). Under an agreement between NASA and ESA, the NASA Orion MPCV for human space exploration missions will be powered by a European Service Module, based on the design and experience of the ATV (Automated Transfer Vehicle). The development and qualification of the European Service Module is managed and implemented by ESA. The ESM prime contractor and system design responsible is Airbus Defence and Space. Thales Alenia Space Italia is responsible for the design and integration of the ESM Structure and MMOD protection system in addition to the Thermal Control System and the Consumable Storage System. The Orion Multi-Purpose Crew Vehicle is a pressurized, crewed spacecraft that transports up to four crew members from the Earths surface to a nearby destination or staging point. Orion then brings the crew members safely back to the Earths surface at the end of the mission. Orion provides all services necessary to support the crew members while on-board for short duration missions (up to 21 days) or until they are transferred to another orbiting habitat. The ESM supports the crew module from launch through separation prior to re-entry by providing: in-space propulsion capability for orbital transfer, attitude control, and high altitude ascent aborts; water and oxygen/nitrogen needed for a habitable environment; and electrical power generation. In addition, it maintains the temperature of the vehicle's systems and components and offers space for unpressurized cargo and scientific payloads. The ESM has been designed for the first 2 Lunar orbit missions, EM-1 (Exploration mission 1) is an un-crewed flight planned around mid-2020, and EM-2, the first crewed flight, is planned in 2022. At the time where the first ESM is about to be weighted, the predicted mass lies slightly above the initial requirement. For future builds, mass reduction of the Service Module has been considered necessary. This is being investigated, together with other design improvements, in order to consolidate the ESM design and increase possible future missions beyond the first two Orion MPCV missions. The mass saving study has introduced new optimised structural concepts, optimisation of the MMOD protection shields, and optimised redesign of parts for manufacturing through AM (Additive Manufacturing).

Description: The interaction of on-axis and o -axis laser discharge in front of a hemisphere cylinder in Mach 2.0 ow is investigated numerically. Details of the physics of the interaction of the laser-induced shock and the heated region with the bow shock and its e ect on drag reduction are included. The energetic eciency of the laser discharge in reducing drag is calculated.

Description: Active flow control (AFC) in the form of sweeping jet (SWJ) excitation and discrete steady jet excitation is used to control the flow separation on an NACA 0015 semispan wing with a deflected, simple-hinged, trailing edge flap. This geometry has been the focus of several recent publications that investigated methods to improve the efficiency of sweeping jet actuators. In the current study, the interaction of the AFC excitation with the separated flowfields present at several flap deflection angles was examined. Previous studies with this model have been limited to a maximum flap deflection angle of 40. The flap deflection range was extended to 60! because systems studies have indicated that a high-lift system with simple-hinged flaps may require larger flap deflections than the Fowler flaps found on most high-lift systems. The results obtained at flap deflection angles of 20, 40, and 60 are presented and compared. Force and moment data, Particle Image Velocimetry (PIV) data, and steady and unsteady surface pressure data are used to describe the flowfield with and without AFC. With a flap deflection of 60, increasing the SWJ actuator momentum at the flap shoulder increased lift due to an increase in circulation but did not completely eliminate the recirculation region above the flap surface. AFC using the discrete steady jet actuators of this study increased lift as well but required more mass flow than the SWJ actuators and had a detrimental effect on lift at the highest mass flow level tested. PIV results showed that the angle between the excitation and the flap surface was not optimal for attaching the separated shear layer.

Description: The FAST-MAC circulation control model was modified to test an array of steady and unsteady actuators at realistic flight Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center. Previous experiments in the FAST-MAC test series used a fullspan tapered slot, and that configuration is used as a baseline for performance and weight flow requirements. The goal of the latest experiment was to reduce the weight flow required to achieve comparable performance established by the baseline FAST-MAC data. Thirty-nine interchangeable actuator cartridges of various designs were mounted into the FAST-MAC model where the exiting jet was directed over a 15% chord simple hinged-flap. These two types of actuators were fabricated using rapid prototype techniques and their design performance was optimized for a transonic cruise configuration having a 0 flap deflection. The steady actuators were found to provide an off-design drag reduction of 5.5%, nearly equaling the drag reduction of the fullspan tapered slot configuration, but with a 69% weight flow reduction. This weight flow savings is similar to the sweeping jet actuators, but with better drag performance.

Description: Various methods for remote recession sensing of PICA have been developed and several seeding methods have been tested. The most recent method involved seeding the ablator with wires fed to the sample from the backside with a defined amount of PICA left towards the upstream front of the sample. This seed method mimics the installation of in-depth thermocouples as they are frequently used in ground testing and flight. Arc-jet tests were conducted in the NASA Langley HYMETS facility at a heat flux of 320 W/sq.cm. The emission of the post-shock layer was observed in spectral resolution from the side along an optical axis perpendicular to the arc-jet flow and from the front, looking at the sample surface from an upstream position. Various metallic seed materials with different melting points were used. In addition to the emission spectroscopy measurements, the samples were monitored during the tests through pyrometry and videography. The time resolved response of the seeded material is described and compared to earlier tests with different seeding methods. The combination of seed materials was found to be critical for the selection of emission signatures characteristic for the material recession which can be isolated in the final emission spectra.

Description: The Molecular Adsorber Coating (MAC) is a sprayable coatings technology that was developed at NASA Goddard Space Flight Center (GSFC). The coating is comprised of highly porous, zeolite materials that help capture outgassed molecular contaminants on spaceflight applications. The adsorptive capabilities of the coating can alleviate molecular contamination concerns on or near sensitive surfaces and instruments within a spacecraft. This paper will discuss the preliminary testing of NASA's MAC technology for use on future missions to Mars. The study involves evaluating the coating's molecular adsorption properties in simulated test conditions, which include the vacuum environment of space and the Martian atmosphere. MAC adsorption testing was performed using a commonly used plasticizer called dioctyl phthalate (DOP) as the test contaminant.

Description: Final document is attached. This paper proposes an enhanced control technique for stationkeeping maneuvers to reduce delta-v costs for the Korea Pathfinder Lunar Orbiter (KPLO). A scheduled circularization control technique exploits patterns in the evolution of the line of apsides and eccentricity to achieve a significant reduction in stationkeeping delta-v costs based on spacecraft requirements. The technique is compared against previous algorithms implemented for maneuver operations of the Lunar Prospector and Lunar Reconnaissance Orbiter (LRO) missions in the USA and KAGUYA in Japan. Through Monte Carlo analysis, the efficacy and robustness of the proposed method are verified, and the technique is shown to meet the operational requirements of KPLO.

Description: Final document not an Abstract attached. The International Space Station (ISS) has been on-orbit for nearly 20 years, and there have been numerous technical challenges along the way from design to assembly to on-orbit anomalies and repairs. The Passive Thermal Control System (PTCS) management team has been a key player in successfully dealing with these challenges. The PTCS team performs thermal analysis in support of design and verification, launch and assembly constraints, integration, sustaining engineering, failure response, and model validation. This analysis is a significant body of work and provides a unique opportunity to compile a wealth of real world engineering and analysis knowledge and the corresponding lessons-learned. The PTCS lessons encompass the full life cycle of flight hardware from design to on-orbit performance and sustaining engineering. These lessons can provide significant insight for new projects and programs. Key areas to be presented include thermal model fidelity, verification methods, analysis uncertainty, and operations support.

Description: The flight focal plane array (FPA) for the Thermal Infrared Sensor 2 (TIRS-2) instrument, to be flown on Landsat 9, was built and characterized at NASA Goddard Space Flight Center (GSFC). The FPA was assembled using GaAs quantum well infrared photodetector (QWIP) arrays from the same lot as the TIRS instrument on Landsat 8. Each QWIP array is hybridized to an Indigo ISC9803 readout integrated circuit (ROIC) with 640 x 512, 25m by 25m pixels. Each QWIP hybrid was tested at the NASA/GSFC Detector Characterization Laboratory (DCL) as a single sensor chip assembly (SCA). The best SCAs in terms of performance were then built up into an FPA consisting of three SCAs, required to provide the necessary 15-degree field of view of the instrument. The FPA was tested to determine if project requirements were being met as a fully assembled unit. The performance of the QWIP SCAs and the fully assembled, NASA flight-qualified FPA will be reviewed.

Description: Direct Field Acoustic Testing (DFAT) offers potential cost and time savings over reverberant chamber acoustic testing of spacecraft. The NASA Multi-Purpose Crew Vehicle (MPCV) Program recently directed a series of acoustic tests on Orion structural test articles comparing DFAT and reverberant testing of the same test article with a view to qualifying DFAT for manned space flight vehicles. The verification process compared four parameters noise level compliance with the one third octave test specification, spatial uniformity of the acoustic field, spatial correlation of the acoustic field and vibration response of vehicle structure, including representative solar array panels. While the results of the verification were encouraging, MPCV Loads and Dynamics engaged Quartus Engineering to investigate whether alternative MIMO random control strategies might improve the spatial uniformity and/or the spatial correlation of the DFAT acoustic field. This paper presents the results of acoustic field simulations of the DFAT test and provides a better understanding of how MIMO random control systems originally developed for vibration and structural durability testing can be expected to perform in DFAT testing.

Description: In order to optimize systems, systems engineers require some sort of measure with which to compare vastly different system components. One such measure is system exergy, or the usable system work. Exergy balance analysis models provide a comparison of different system configurations, allowing systems engineers to compare different systems configuration options. This paper presents the exergy efficiency of several Mars transportation system configurations, using data on the interplanetary trajectory, engine performance, and vehicle mass. The importance of the starting and final parking orbits is addressed in the analysis, as well as intermediate hyperbolic escape and entry orbits within Earth and Mars' spheres of influence (SOIs). Propulsion systems analyzed include low-enriched uranium (LEU) nuclear thermal propulsion (NTP), high-enriched uranium (HEU) NTP, LEU methane (CH4) NTP, and liquid oxygen (LOX)/liquid hydrogen (LH2) chemical propulsion.

Description: The Near Earth Asteroid (NEA) [1] Scout is a deep space CubeSat designed to use an 86 m2 solar sail to navigate to a near earth asteroid called VG 1991. The solar sail deployment mechanism aboard NEA Scout has gone through numerous design cycles and ground tests since its conception in 2014. An engineering development unit (EDU) was constructed in the spring of 2016 and since then, the NEA Scout team has completed numerous ground deployments aiming to mature the deployment system and the ground test methods used to validate that system. Testing a large, non-rigid gossamer system in 1G environments has presented its difficulties to numerous solar sailing programs before, but NEA Scout's size, sail configuration, and budget has led the team to develop new deployment techniques and uncover new practices while improving their test methods. The program has planned and completed 5 separate full scale sail deployments to date, with a flight sail deployment test scheduled for FY18. The paper entitled "Design and Development of NEA Scout Solar Sail Deployer Mechanism" [2] was presented at the 43rd Aerospace Mechanisms Symposia. Since then, the system has matured and completed ascent vent, random vibration, boom deployment and sail deployment tests. This paper will discuss the lessons learned and advancements made while working on solar sail deployment testing and mechanical redesign cycles.

Description: For the last 5 years, NASA Goddard has been investigating Distributed Spacecraft Missions (DSM) system architectures, surveying past, current and potential mission concepts, developing several taxonomies and identifying some key technologies that will enable future DSM mission design, development, operations and management. This paper summarizes this Initiative and the talk will provide details about specific Goddard DSM projects that are currently underway and that are relevant to future Earth Science missions.

Description: This paper presents an overview of the development and qualification test campaign for the primary structure of the European Service Module of ORION, the NASA spacecraft which will serve the future human exploration missions to the Moon, Mars and beyond. Under an agreement between NASA and ESA, the ORION will be powered by a European Service Module (ESM), providing also water and oxygen for astronauts' life sustainability. The development and qualification of the European Service Module (ESM) is under ESA responsibility with Airbus Defense and Space as the prime contractor. Thales Alenia Space Italia is responsible for design development, manufacturing, assembly and qualification of the Structure subsystem. The European Service Module, installed onto the launch adapter, shall support the crew module with its adapter and a launch abort system. It shall sustain: - A combination of global and local launch loads during lift off and ascent phases, - On orbit loads induced by engine firing for orbital transfers and attitude control. The ESM structure is based on a core made of Composite Fiber Reinforced Polymer (CFRP) sandwich panels complemented by aluminum alloy platforms, longerons and secondary structures. A development campaign has been implemented in order to define and validate composite parts' strength allowable values for design: coupon tests at material level, test at component level up to breadboards tests performed on main structural components (composite to metallic joints, and at panels' discontinuities). An incremental approach as defined in [1] has been followed. A qualification static test campaign at primary structure assembly level has been implemented in order to validate the design against static stiffness and ultimate strength as well as to correlate the structural Finite Element Model (FEM) used for sizing and confirm the margins of safety. The tests have been performed successfully by Thales Alenia Space Italia (TAS-I) on two flight representative structural models (STA1, STA2), in Turin facilities (Italy) between August 2015 and March 2017, with engineering support of technical representatives from Airbus, ESA, NASA and LMCO. The main development and qualification test activities and associated results are presented and discussed in the paper

Description: As NASA looks towards human missions to Mars, an effort has started to advance the technology of a Mars in situ resource utilization (ISRU) Propellant Production Plant to a flight demonstration. This paper will present a design study of the Sabatier subsystem. The Sabatier subsystem receives carbon dioxide, CO2, and hydrogen, H2, and converts them to methane, CH4, and water, H2O. The subsystem includes the Sabatier reactor, condenser, thermal management, and a recycling system (if required). This design study will look at how the choice of reactor thermal management, number of reactors, and recycling system affect the performance of the overall Sabatier system. Different schemes from the literature involving single or cascading reactors will be investigated to see if any provide distinct advantages for a Mars propellant production plant.

Description: No abstract available

Description: No abstract available

Description: Propellant slosh was analyzed for both the oxidizer and the fuel for the Europa Clipper propulsion system. Slosh was examined for various fill fractions for cases where acceleration was on the order of magnitude of 10(exp -2) m/sq. s using the computational fluid dynamics software package STAR-CCM+ and at various fill fractions for cases where acceleration was on the order of magnitude of 10(exp -5) m/sq. s using Surface Evolver. Equivalent mechanical model parameters were derived from the CFD data using MATLAB for both the higher and the lower acceleration slosh cases. These parameters were plotted and can be used to interpolate mechanical model parameters at fill fractions not analyzed by CFD or Surface Evolver.

Description: System engineering of launch vehicles and spacecraft is a challenging and complex undertaking. There are many diverse systems which must be integrated and balanced to produce an effective design. This involves a multiplicity of individual engineering relationships that are difficult to integrate and even more difficult to define in a best balance. Integration efforts involve many different approaches, from process management to mass balance. But these approaches either do not directly address the launch vehicle or spacecraft performance or require many adjustments to be made to discover a balance. The system integrating physics, derived from the fundamental physics of the system, is the key to identifying a fully integrated system performance measure. Launch vehicles and spacecraft are thermodynamic systems with performance defined by thermodynamic properties. Thus, thermodynamic exergy, which integrates all of the systems thermodynamic properties, provides the system integrating relationships. This provides a basis for determining the most efficient design from among many different configuration options and for guiding the design activities from an integrated system level. This paper explores the current physics relationships used in launch vehicle system design and demonstrates that thermodynamic exergy provides a more explicit and complete approach to system integration.

Description: Atmospheric probes have been successfully flown to planets and moons in the solar system to conduct in situ measurements. They include the Pioneer Venus multi-probes, the Galileo Jupiter probe, and Huygens probe. Probe mission concepts to five destinations, including Venus, Jupiter, Saturn, Uranus, and Neptune, have all utilized similar-shaped aeroshells and concept of operations, namely a 45-degree sphere cone shape with high density heatshield material and parachute system for extracting the descent vehicle from the aeroshell. Each concept designed its probe to meet specific mission requirements and to optimize mass, volume, and cost. At the 2017 International Planetary Probe Workshop (IPPW), NASA Headquarters postulated that a common aeroshell design could be used successfully for multiple destinations and missions. This "common probe" design could even be assembled with multiple copies, properly stored, and made available for future NASA missions, potentially realizing savings in cost and schedule and reducing the risk of losing technologies and skills difficult to sustain over decades. Thus the NASA Planetary Science Division funded a study to investigate whether a common probe design could meet most, if not all, mission needs to the five planetary destinations with extreme entry environments. The Common Probe study involved four NASA Centers and addressed these issues, including constraints and inefficiencies that occur in specifying a common design. Study methodology: First, a notional payload of instruments for each destination was defined based on priority measurements from the Planetary Science Decadal Survey. Steep and shallow entry flight path angles (EFPA) were defined for each planet based on qualification and operational g-load limits for current, state-of-the-art instruments. Interplanetary trajectories were then identified for a bounding range of EFPA. Next, 3-degrees-of-freedom simulations for entry trajectories were run using the entry state vectors from the interplanetary trajectories. Aeroheating correlations were used to generate stagnation point convective and radiative heat flux profiles for several aeroshell shapes and entry masses. High fidelity thermal response models for various Thermal Protection System (TPS) materials were used to size stagnation-point thicknesses, with margins based on previous studies. Backshell TPS masses were assumed based on scaled heat fluxes from the heatshield and also from previous mission concepts. Presentation: We will present an overview of the study scope, highlights of the trade studies and design driver analyses, and the final recommendations of a common probe design and assembly. We will also indicate limitations that the common probe design may have for the different destinations. Finally, recommended qualification approaches for missions will be presented.

Description: NASA's Orion exploration spacecraft will fly more demanding mission profiles than previous NASA human flight spacecraft. Missions currently under development are destined for cislunar space. The EM-1 mission will fly unmanned to a Distant Retrograde Orbit (DRO) around the Moon. EM-2 will fly astronauts on a mission to the lunar vicinity. To fly these missions, Orion requires powered flight guidance that is more sophisticated than the orbital guidance flown on Apollo and the Space Shuttle. Orion's powered flight guidance software contains five burn guidance options. These five options are integrated into an architecture based on a proven shuttle heritage design, with a simple closed-loop guidance strategy. The architecture provides modularity, simplicity, versatility, and adaptability to future, yet-to-be-defined, exploration mission profiles. This paper provides a summary of the executive guidance architecture and details the five burn options to support both the nominal and abort profiles for the EM-1 and EM-2 missions.

Description: This poster provides an overview of the requirements, design, development and testing of the 3D (Three Dimensional) Woven TPS (Thermal Protection System) being developed under NASA's Heatshield for Extreme Entry Environment Technology (HEEET) project. Under this current program, NASA is working to develop a TPS capable of surviving entry into Saturn. A primary goal of the project is to build and test an Engineering Test Unit (ETU) to establish a Technical Readiness Level (TRL) of 6 for this technology by 2017.

Description: No abstract available

Description: The aim of the Distributed Attitude Control and Maneuvering for Deep Space SmallSats project is to advance a multi-purpose, deep space mission-enabling technology for low-power attitude and thermal control of small satellites to a flight demonstration technology readiness level (TRL). The film-evaporation microelectromechanical systems tunable array (FEMTA) small satellite technology combines innovative microelectromechanical systems (MEMS) microfabrication and microscale effects in fluid surface tension to produce a thermally actuated capillary valve. Using water as the propellant, the FEMTA thruster can generate finely controllable thrust at a thrust to power ratio of about 200 microNewton per Watt (W).

Description: The Starling series of demonstration missions will test technologies required to achieve affordable, distributed spacecraft ("swarm") missions that: are scalable to at least 100 spacecraft for applications that include synchronized multipoint measurements; involve closely coordinated ensembles of two or more spacecraft operating as a single unit for interferometric, synthetic aperture, or similar sensor architectures; or use autonomous or semi-autonomous operation of multiple spacecraft functioning as a unit to achieve science or other mission objectives with low-cost small spacecraft.Starling1 will focus on developing technologies that enable scalability and deep space application. The mission goals include the demonstration of a Mobile Ad-hoc NETwork (MANET) through an in-space communication experiment, vision based relative navigation through the Starling Formation-flying Optical eXperiment (StarFOX), and demonstration of autonomous spacecraft reconfiguration using technologies developed by the Distributed System Autonomy (DSA) project.

Description: Atmospheric probes have been successfully flown to planets and moons in the solar system to conduct in-situ measurements. They include the Pioneer Venus multi-probes, the Galileo Jupiter probe, and Huygens probe. Probe mission concepts to five destinations, including Venus, Jupiter, Saturn, Uranus, and Neptune, have all utilized similar-shaped aeroshells and concept of operations, namely a 45 deg sphere cone shape with high density heatshield material and parachute system for extracting the descent vehicle from the aeroshell. Each concept designed its probe to meet specific mission requirements and to optimize mass, volume, and cost. At the 2017 IPPW, NASA Headquarters postulated that a common aero-shell design could be used successfully for multiple destinations and missions. This "common probe" design could even be assembled with multiple copies, properly stored, and made available for future NASA missions, potentially realizing savings in cost and schedule and reducing the risk of losing technologies and skills difficult to sustain over decades. Thus the NASA Planetary Science Division funded a study to investigate whether a common probe design could meet most, if not all, mission needs to the five planetary destinations with extreme entry environments. The Common Probe study involved four NASA Centers and addressed these issues, including constraints and inefficiencies that occur in specifying a common design.Study methodology: First, a notional payload of instruments for each destination was defined based on priority measurements from the Planetary Science Decadal Survey. Steep and shallow entry flight path angles (EFPA) were defined for each planet based on qualification and operational g-load limits for current, state-of-the-art instruments. Interplanetary trajectories were then identified for a bounding range of EFPA. Next, 3-DoF simulations for entry trajectories were run using the entry state vectors from the interplanetary trajectories. Aeroheating correlations were used to generate stagnation point convective and radiative heat flux profiles for several aeroshell shapes and entry masses. High fidelity thermal response models for various TPS materials were used to size stagnation point thicknesses, with margins based on previous studies. Backshell TPS masses were assumed based on scaled heat fluxes from the heatshield and also from previous mission concepts.

Description: This paper presents an overview of the Second European Service Module (ESM-2), the second in a series of European Service Modules produced as part of the Barter agreement between NASA and ESA for the Orion Program. The European Industrial consortium is led by the ESA prime contractor Airbus Defence and Space in Bremen. ESA and Airbus signed the ESM-2 contract on 16 February 2017, for this key element of the Orion Exploration Mission 2 (EM-2). EM-2 is the first crewed mission for Orion and will take astronauts farther into the solar system than humanity has ever travelled. EM-2 will also be a historic mission for Europe, as the ESM-2 will be the first European spacecraft to be part of a human transportation system carrying humans beyond low Earth orbit. ESM-2 is mainly a recurring production following ESM-1. Nevertheless, there are a number of important changes being implemented, for example, to incorporate upgrades to further enhance safety and reliability. The challenging delivery schedule for ESM-2 has driven the need to commence manufacturing prior to completion of the qualification on ESM-1. In addition, some requirement deviations and non-compliances approved for ESM-1 have resulted in modifications for ESM-2. In order to manage the competing constraints effectively, the ESM-2 Team has put in place a number of novel approaches to manage schedule, risk, and technical changes. Airbus has set up multi-functional teams according to an approach known as "Major Spacecraft Deliveries" consisting of quality assurance, engineering and procurement. The risk of starting manufacturing prior to qualification is managed through a special risk share agreement. This agreement necessitates rigorous risk reviews across the board for all manufacturing, assembly, integration and test milestones. The ESM-2 changes are managed by Configuration Management, but Airbus has also introduced the Technical Baseline Matrix to provide a transparent top-level overview of the changes from ESM-1 to ESM-2. The tool provides the basis for ESM-2 design and development needs, decisions, as well as the input for the Orion EM-2 Critical Design Review (CDR). The main technical evolutions, status of the production and the novel management approaches for ESM-2 are presented and discussed in the paper.

Description: Following a very successful year of manufacturing, assembly and testing in factories located around the globe, NASA and ESA are preparing to deliver the major Exploration Mission-1 (EM-1) Orion flight elements, including the Crew Module, ESA Service Module and Launch Abort System. This international effort to design and develop a deep space exploration capable human spacecraft is rapidly transitioning from the design, development and test phase to the early test flight and production phase. Two major flight tests, an Ascent Abort test and EM-1, Orion's first flight onboard NASA's new heavy lift Space Launch System, are planned for the near future. Further, Orion will play a crucial role in the ambitious new Deep Space Gateway human exploration Program. This paper gives a short overview of the system and subsystem configuration of the Orion spacecraft, including NASA and ESA contributions, a status of EM-1, AA-2 and EM-2 spacecraft production, and a look at Orion's role in the construction and operation of the Deep Space Gateway. The paper will also address the innovative international cooperation methods being employed to conduct Orion and Service Module integration.

Description: A regeneratively-cooled nozzle for liquid rocket engine applications is a significant cost of the overall engine due to the complexities of manufacturing a large thin-walled structure that must operate in extreme temperature and pressure environments. The National Aeronautics and Space Administration (NASA) has been investigating and advancing methods for fabrication of liquid rocket engine channel wall nozzles to realize further cost and schedule improvements over traditional techniques. The methods being evaluated are targeting increased scale required for current NASA and commercial space programs. Several advanced rapid fabrication methods are being investigated for forming of the inner liner, producing the coolant channels, closeout of the coolant channels, and fabrication of the manifolds. NASA's Marshall Space Flight Center (MSFC) has completed process development and subscale hot-fire testing of a series of these advanced fabrication channel wall nozzle technologies to gather performance data in a relevant environment. The primary fabrication technique being discussed in this paper is Laser Wire Direct Closeout (LWDC). This process has been developed to significantly reduce the time required for closeouts of regeneratively-cooled slotted liners. It allows for channel closeout to be formed in place in addition to the structural jacket without the need for channel fillers or complex tooling. Additional technologies were also tested as part of this program including water jet milling and arc-based additive manufacturing deposition. Each nozzle included different fabrication features, materials, and methods to demonstrate durability in a hot-fire environment. The results of design, fabrication, and hot-fire testing are discussed in this paper.

Description: In liquid propellant rocket engines, spark igniters are often used indirectly to light preburners, gas generators, and main chambers [1]. Attraction for spark igniters is strongly influenced by their ability for repeatable engine starts and high reliability. In the case of spark igniters, however, ignition is reliant upon an ignitable mixture passing near the spark tip very early in the engine start transient, prior to pressure quenching of the spark. While direct ignition of rocket engine combustion chambers is possible and has been successfully implemented in engines such as RL-10, the development time can be significant since ignition requires precise and repeatable control of the propellant mixture ratio within the very small volume and short duration of the spark plasma. Generally, the preferred method of implementing spark igniters within rocket engines - especially larger engines, is to design a smaller "augmented spark igniter" pre-chamber in which propellant injection and mixture ratio near the spark plasma can be controlled independent of the engine injector. The resultant combustion products within the small pre-chamber are directed into the larger engine chamber via a torch tube. An augmented spark igniter is advantageous because the output torch flame that is much larger and more energetic than a discrete train of small spark plasmas.

Description: The James Webb Space Telescope Primary Mirror Segment Assemblies (PMSAs) and Secondary Mirror Assembly (SMA) were cleaned at the Johnson Space Center (JSC) in January 2018. In order to quantify the effectiveness of the cleaning, the same cleaning process was performed on the PMSA and SMA traveling witness wafers. These wafers have accompanied their respective mirror segments from their arrival at the Goddard Space Flight Center, through transport to JSC, and ultimately their exposure in Chamber A for cryogenic testing. The traveling wafers were analyzed using an Image Analysis automated microscope both prior to and after the cleaning. The resulting data showed that the PMSA wafers' Percent Area Coverage (PAC) reduced by 83.5% on average, from 0.1524 PAC to 0.0251 PAC. The SMA wafer's PAC decreased by 97.2%, from 0.1194 PAC to 0.0034 PAC. Further analysis of the particle size bins was completed in order to calculate their particle distribution slopes. The slope of the PMSA wafers increased by 0.025 on average, and the SMA wafer slope increased by 0.066. This indicates that the ratio of large to small particles slightly increased after the cleaning across all mirror segments. Visual inspections of the wafers and the flight PMSAs and SMA showed considerable and comparable particulate coverage improvements, thus leading to the conclusion that the average PAC on the PMSAs and SMA improved by the same factor as their respective wafers.

Description: No abstract available

Description: Presently, most CubeSat components and buses are generally not appropriate for missions where significant or indeterminate risk of failure is unacceptable. This has precluded their use in many cases where their attributes could otherwise enable or enhance mission objectives. However, in the future, CubeSats and SmallSats, which deviate from CubeSat form factors but often incorporate CubeSat components and subsystems, will address challenges that many presently consider to be beyond the platform's capabilities. This growing potential utility, combined with the limited volume of successful CubeSat flight heritage, is driving an interagency effort to improve small satellite mission confidence.

Description: Recent introduction of Coaxial Thermocouple type calorimeters into the NASA Ames arc jet facilities has inspired an analysis of 2D conduction effects internal to this type of calorimeter. The 1D finite slab inverse analysis (which is typically used to deduce the heat transfer to the calorimeter) relies on the assumption that lateral conduction (i.e., 2D effects) is negligible. Most calorimeter bodies have a spherical nose, which in itself is a violation of the 1D finite slab analysis assumption. Secondly most calorimeters experience a variation in heating across the face of the body which is also a violation of the 1D finite slab analysis assumption. It turns out that these two effects tend to cancel each other to some extent. This paper shows the extent to which error exists in the analysis of the Coaxial Thermocouple type calorimeters, and also offers analysis strategies for reducing the errors.

Description: The Mars2020 entry vehicle is currently being developed by NASA to safely land its next rover on the Martian surface in 2021. The vehicle will be protected from entry aeroheating using three different TPS materials: PICA tiles on the forebody, SLA-561V on the backshell and Acusil-II on the parachute close-out cone (PCC) and its backshell interface plate (BIP). Mars2020's entry vehicle and TPS design is identical to the Mars Science Laboratory, NASA's last Mars lander; therefore, the purpose of this study is to assess the adequacy of the existing TPS design and thickness for Mars2020 predicted environments. This study focuses on sizing and margin assessment of Acusil-II TPS on the PCC and BIP. The methodology and analysis techniques that were used for assessing thermal margins are reviewed. Analysis assumptions and limitations are discussed in detail. Thermal sizing is performed at different locations and results are presented.

Description: No abstract available

Description: Lynx is an X-Ray telescope large-mission concept for consideration in NASA's 2020 Astrophysics Decadal Survey. A conceptual structural design is evolving that leverages the success and lessons learned from Chandra and that takes into account unique needs of Lynx. Space optics systems require extreme stability. Any motion in-service (thermal effects, structural dynamics, etc.) impacts performance. An initial analysis was performed to predict the first-cut dynamic responses, jitter, at two selected points on the Lynx observatory. One point is on the Lynx X-ray Mirror Assembly (LMA) and the other, on the focal plane Integrated Science Instrument Module (ISIM). Relative motion between these two points was predicted along with vibration spectra. This information will be used in upcoming analyses of the LMA and the ISIM.

Description: No abstract available

Description: The James Webb Space Telescope (JWST), set to launch in early 2019, is currently undergoing a series of system-level environmental tests to verify its workmanship and end-to-end functionality. As part of this series, the Optical Telescope Element and Integrated Science Instrument Module (OTIS) Cryo-Vacuum (CV) test, the most complex cryogenic test executed to date by NASA, has recently been completed at the Johnson Space Center's Chamber A facility. The OTIS CV test was intended as a comprehensive test of the integrated instrument and telescope systems to fully understand its optical, structural, and thermal performance within its intended flight environment. Due to its complexity, extensive pre-test planning was required to ensure payload safety and compliance with all limits and constraints. A system-level pre-test thermal model was constructed which fully captured the behavior of the payload, ground support equipment, and surrounding test chamber. This thermal model simulated both the transient cooldown to and warmup from a 20K flight-like environment, as well as predicted the payload performance at cryo-stable conditions. The current work is a preliminary assessment of thermal model performance against actual payload response during the OTIS CV test. It examines both the benefits and shortcomings of assumptions made pre-test to simplify model execution when compared against test data. It explores in detail the role of temperature-dependent emissivities during transition to cryogenic temperatures, as well as the impact that model geometry simplifications have on tracking of critical hardware limits and constraints. This work concludes with a list of recommendations to improve the accuracy of thermal modeling for future large cryogenic tests. It is hoped that the insight gained from the OTIS CV test thermal modeling will benefit planning and execution for upcoming cryogenic missions.

Description: The James Webb Space Telescope (JWST), set to launch in mid-2020, is currently undergoing a series of system-level environmental tests to verify its workmanship and end-to-end functionality. As part of this series, the Optical Telescope Element and Integrated Science Instrument Module (OTIS) Cryo-Vacuum (CV) test, the most complex cryogenic test executed to date by NASA, has recently been completed at the Johnson Space Center's Chamber A facility. The OTIS CV test was intended as a comprehensive test of the integrated instrument and telescope systems to fully understand its optical, structural, and thermal performance within its intended flight environment. Due to its complexity, extensive pre-test planning was required to ensure payload safety and compliance with all limits and constraints. A system-level pre-test thermal model was constructed which fully captured the behavior of the payload, ground support equipment, and surrounding test chamber. This thermal model simulated both the transient cooldown to and warmup from a 20 K flight-like environment, as well as predicted the payload performance at cryo-stable conditions. The current work is an assessment of thermal model pre-test prediction performance against actual payload response during the OTIS CV test. Overall, the thermal model performed exceedingly well at predicting schedule and payload response. Looking in depth, this work examines both the benefits and shortcomings of assumptions made pre-test to simplify model execution when compared against test data. It explores in detail the role of temperature-dependent emissivities during transition to cryogenic temperatures, as well as the impact that model geometry simplifications have on tracking of critical hardware limits and constraints. This work concludes with a list of recommendations to improve the accuracy of thermal modeling for future large cryogenic tests. The insight gained from the OTIS CV test thermal modeling will benefit planning and execution for upcoming cryogenic missions.

Description: No abstract available

Description: The Transiting Exoplanet Survey Satellite (TESS) is a NASA Explorer mission. The TESS Observatory is scheduled to launch on Falcon 9 in April 2018. This presentation covers the process used to define and update design limit loads for the observatory, instrument, and components throughout the life of the program. The limit loads that drove the need for a SoftRide isolation system are highlighted. The testing performed to qualify the observatory for launch loads at the instrument and observatory level is also detailed. In addition, exchanges with the launch vehicle provider in terms of loads predictions and hardware for test are discussed along with the associated issues encountered and lessons learned. The loads development and verification success on TESS was a team effort. Orbital ATK is the spacecraft provider, NASA GSFC provides project management and technical oversight, the instrument is managed by MIT Kavli Institute and the instrument cameras are built and tested by MIT Lincoln Laboratory. Since the instrument was designed in parallel with the spacecraft, the instrument design limit loads were developed in partnership with NASA and the instrument team. The three teams collaborated on a regular basis starting in the early design phase and continuing through observatory level testing.