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At least 1,099 records · Page 61
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Fluid-Structure Interaction Simulations Of Supersonic Parachute Inflation: Model Sensitivities

Fluid-structure interaction simulations have the potential to accelerate the certification process of parachute systems for interplanetary robotic exploration missions while helping to mitigate risks by covering a wider range of scenarios than can be flight tested. The Launch, Ascent, and Vehicle Aerodynamics team is working toward actualizing this potential by further developing the capability to perform such fluid-structure interaction simulations and validating it with the best supersonic parachute flight test data available. We perform several simulations of the Advanced Supersonic Parachute Inflation Research Experiments’ third flight test (SR03) where we independently vary several parameters to determine their sensitivity on the predicted drag. We investigate sensitivity of the drag to the added mass of the confluence fitting, to the effective porosity, to how the gore seams are modeled, to viscous effects, and to deceleration. Modeling the added mass of the confluence fitting has virtually no impact on the drag, but is worthwhile to include to better match the overall system’s mass, which is important for the simulations with deceleration. Parachute drag is only marginally sensitive to porosity and viscous effects. Maximum parachute drag is most sensitive to how the gore seams are modeled, causing a nearly 20% drop when they are modeled as stiffer radials, and to deceleration, causing an additional 10% reduction. Results show improved agreement with flight test measurements when the FSI model is closest to the as-built configuration and physics experienced in flight: with the confluence fitting, viscous effects, the gore seams modeled as stiff radials, and including deceleration.

ESM↗
NASA NTRSDate not supplied

ISS External Microorganisms: Collecting Planetary Protection Samples During Extravehicular Activity

We have developed, tested, and flown a caddy capable of collecting aseptic samples from external surfaces of the ISS (International Space Station). The sampling caddy is certified for use during US EVA (extra vehicular activity) and was launched to ISS in the summer of 2023. We are scheduled to collect samples from 6 locations outside ISS during an EVA in May of 2024. We will freeze these samples at -80°C on orbit and return them to Earth. We will then extract and sequence any DNA collected during the EVA using next generation sequencing technologies to characterize the community composition and function of each sample. Measuring the type and quantity of microbes present on the exterior of ISS will allow us to address knowledge gap 2B, “Acceptable levels of microbial/organic releases from humans and support systems” described in a 2019 COSPAR report. Collecting data about microbial release from current crewed vehicles will inform requirements for acceptable leak rates for future crewed missions to Mars. The sampling kit consists of eight commercially available, sterile, DNA free, macrofoam swabs ( 23 mm. diameter ) installed in custom aluminum end effectors. Each end effector is housed in and individual aluminum canister. Each canister contains a 0.2 μm Teflon filter to allow the interior volume to accommodate pressure changes without permitting microbial contaminants to enter the sterile interior volume. A handle repurposed from the space shuttle tile repair kit is used to remove the end effector from the sample canister, collect a sample by swabbing a surface and then replace the end effector in its canister. The canisters and end effectors were cleaned and assembled on Earth. Prior to installing the sterile swab the canisters and end effectors were sterilized in an autoclave at 134°C, 215 kPA, for 7 min.. The final assembly occurred in a sterilized class II biosafety cabinet. We will collect six samples from the 1) airlock vestibule, 2) airlock thermal cover, 3) a gap in the micrometeorite shielding near the airlock, 4) a handrail near the airlock, 5) the CDRA (Carbon Dioxide Removal Assembly) vent, and 6) the VES (Vacuum Exhaust System) vent. The remaining two swabs will be reserved as controls. One swab will be exposed during the EVA without touching any surfaces to act as a blank. The final swab will remain sealed until the entire sampling kit is returned to earth. Based on previously published results from the Russian segment, we hypothesize that there will be detectable microbes at some or all of these locations. Ground-based testing of this sampling caddy confirms that the swabs remain sterile as the canisters transition in and out of vacuum. We were able to retrieve, viable bacterial and fungal cells as well as DNA from samples collected from US space suits during vacuum chamber tests lasting as long as seven hours. Based on these results and feedback from the test subjects the sampling caddy was modified to improve ergonomics and meet US EVA safety requirements. Bayonet probes were added to the sides of the sample kit as alternate mounting points. Additional locking features were added to the end effector and the filter stack to prevent inadvertent release during use. The opening mechanism was changed from one where the end effector was rocked laterally to defeat a ball detent to a twist-to-open threaded closure for similar reasons. Demonstrating, this sampling caddy’s effectiveness during a US EVA will allow us to address knowledge gaps identified in COSPAR reports and begin to define planetary protection requirements for life support systems on crewed missions to mars. This kit could also be used to collect contamination control samples during Artemis missions to verify requirements and could be easily modified for robotic sample collection.

Planetary Protection↗
NASA NTRSDate not supplied

Validation of Computational Fluid Dynamics Model for Pressurized Expulsion of Liquid Methane using Non-Condensable Pressurant

An important aspect of cryogenic fluid management is the design of the pressurization systems used to maintain and regulate the tank pressure as propellant is expelled from the tanks. To design these systems efficiently and effectively it is important to be able to predict the amount pressurant gas needed and how the pressurant gas affects the propellant as it is added in the tank. Computational fluid dynamics (CFD) modeling is a common tool that is used to help with these predictions, so it is crucial to have CFD models that are validated by comparing numerical results to experimental data. In the 1970’s extensive testing was done at NASA Plum Brook K-Site Facility exploring the pressurant gas requirements needed to maintain tank pressure while liquid methane is discharged from a spherical tank. These tests explored using different pressurant gases (methane, helium, hydrogen, and nitrogen) under static and sloshing conditions. The data gathered from the experiments includes the gas and wall temperatures at the end of the expulsion, the gas concentrations, and the total amount of pressurant added. This study presents a CFD model that accurately predicts similar results to the experimental data specifically for the tests in which helium, a non-condensable gas, was used. This comparison anchors the commercial CFD code Flow3D model and gives confidence for future use in design and analysis of pressurization systems for cryogenic propellant tanks

computational fluid dynamics↗
NASA NTRSDate not supplied

Preliminary Application of Formal Verification to An Autonomy Architecture for Unmanned Aircraft

There is a desire to design autonomous systems in such a way that capabilities can be easily added or re-combined to produce new behaviors while preserving their safety properties. ICAROUS, a prototype software architecture for building safety-centric autonomous unmanned aircraft applications, is designed to support this type of extensibility and re-configurability. In ICAROUS, core capabilities are implemented as individual soft- ware services, so that enabling access to new capabilities simply requires adding new services. To make use of these capabilities, ICAROUS includes a specialized service that provides a general framework for config- uring the relative priorities, conditions, and rules that govern how different modules should be engaged and disengaged during flight. The inherent complexity of coordinating multiple modules under changing conditions makes it difficult to determine whether a particular configuration could have erroneous behaviors in certain circumstances. A robust set of integration tests can help discover errors, but testing can only realistically cover a relatively small proportion of total system behaviors. Developing good tests and interpreting the results to pinpoint the cause of errors when they arise can also be very time-consuming. To supplement testing, formal methods can be used to model and analyze complex systems, achieving better coverage and simplifying the process of finding, understanding, and fixing errors. To demonstrate these benefits, this paper explores the ap- plication of formal methods to ICAROUS. In particular, the Spin model checker is used to specify requirements for and model portions of the system, then verify whether the model satisfies the requirements and find and fix errors when it does not.

Formal Methods↗
NASA NTRSDate not supplied

Common Cause Failures Dominate and Defeat Redundancy

Common cause failures occur when several malfunctions are produced by a single event or process. They are especially damaging when they eliminate an entire set of redundant systems and disable their intended function. Redundancy is used when the individual system failure probability is unacceptably high. Redundancy can improve the overall system failure probability if the failures are independent, but the reliability gain is limited if there are dependent failures having a common cause. No amount of redundancy can reduce the total failure probability below the common cause failure probability. Common cause failures defeat redundancy. Systems with high reliability requirements often use extensive redundancy. These highly redundant systems rarely fail unless all the redundant components providing a particular function fail. Complete failures of such highly redundant systems are then usually common cause failures. Common cause failures are prevalent in highly redundant, high reliability systems. Common cause failures dominate redundancy. Redundant systems may fail due to specification, design, manufacturing, operations, or maintenance problems that disable all the identical redundant systems. Common cause failures typically account for one tenth of all failures. If the failure probability is relatively low and common cause failures are significant, adding more than two or three redundant identical units usually gives little added reliability improvement. Common cause failures can be reduced by using diverse components with different technologies and manufacturers, by separating and shielding subsystems, and by avoiding shared control, power, or location. External events and shared vulnerabilities may still cause common cause failures.

common cause failures↗
NASA NTRSDate not supplied

Common Cause Failures Dominate and Defeat Redundancy

Common cause failures occur when several malfunctions are produced by a single event or process. They are especially damaging when they eliminate an entire set of redundant systems and disable their intended function. Redundancy is used when the individual system failure probability is unacceptably high. Redundancy can improve the overall system failure probability if the failures are independent, but the reliability gain is limited if there are dependent failures having a common cause. No amount of redundancy can reduce the total failure probability below the common cause failure probability. Common cause failures defeat redundancy. Systems with high reliability requirements often use extensive redundancy. These highly redundant systems rarely fail unless all the redundant components providing a particular function fail. Complete failures of such highly redundant systems are then usually common cause failures. Common cause failures are prevalent in highly redundant, high reliability systems. Common cause failures dominate redundancy. Redundant systems may fail due to specification, design, manufacturing, operations, or maintenance problems that disable all the identical redundant systems. Common cause failures typically account for one tenth of all failures. If the failure probability is relatively low and common cause failures are significant, adding more than two or three redundant identical units usually gives little added reliability improvement. Common cause failures can be reduced by using diverse components with different technologies and manufacturers, by separating and shielding subsystems, and by avoiding shared control, power, or location. External events and shared vulnerabilities may still cause common cause failures.

common cause failures↗
NASA NTRS2024

Generating Dominating Sets Using Locally Defined Centrality Measures

The dominating set problem has many practical applications but is well-known to be NP-hard. Therefore, there is a need for efficient heuristic algorithms, especially in applications such as ad hoc wireless networks. Most distributed algorithms proposed in the literature assume that each node has knowledge of the network structure. We propose a distributed heuristic algorithm that uses two rounds of communication, and where each node has only local information, both in terms of network structure and dominating set assignment. First, each node calculates a local centrality measure to determine whether it is part of the dominating set D. The second round guarantees D is a dominating set by adding any non-dominated nodes. We compare several centrality measures and show that the Shapley centrality, derived from the Shapley value in game theory, is theoretically motivated and performs well in practice on several synthetic and real-world networks.

Network↗
NASA NTRSDate not supplied

Impact Assessment of All-Sky TROPICS Microwave Observations on the NASA GEOS Analyses and Forecasts and Progress to Use the Data in the JEDI-GEOS Analysis System

The NASA Time-Resolved Observations of Precipitation structure and storm Intensity with a Constellation of Smallsats (TROPICS) mission is a constellation of small satellites carrying state-of-art microwave temperature and humidity sounders with 12 channels between 91 GHz and 205 GHz frequency. Including TROPICS-pathfinder, launched on 30 June 2021, five TROPICS CubeSats operate and provide temperature and humidity data to NWP and atmospheric retrieval communities. This study is dedicated to evaluating the impact of the TROPICS satellite constellation microwave observations in numerical weather prediction (NWP) using the NASA Goddard Earth Observing System (GEOS). The TROPICS-01 (TROPICS-Pathfinder), TROPICS-03, TROPICS-05, and TROPICS-06 data in all-sky conditions over the ocean during the period of 25 July 2023 and 6 September 2023 are used for assessing forecast impacts on global NWP analysis and five-day forecasts. A series of experiments are carried out to measure the benefits of assimilating observations from only temperature sounders, water vapor sounders, and both sounders. Statistical analysis of the Observing System Experiments (OSEs) results has shown incremental improvements in global model forecast skills for critical geophysical parameters, including temperature, winds, and geopotential heights. The results demonstrate the potential of the TROPICS-like data to positively impact NWP by adding new information to the current observation and forecast system. In another set of experiments, the TROPICS-03, TROPICS-05, and TROPICS-06 data sets are added to the TROPICS-01 one by one to evaluate the impacts of increasing the revisit rate of TROPICS satellite measurements on NWP analysis for a tropical cyclone’s dynamical and microphysical structures. This study offers important insights into the capabilities of a new generation of small satellite microwave radiometers based on emerging technologies, including their unique measurements at 118 GHz and 205 GHz that are not available in traditional operational microwave sounders. Finally, the efforts to implement these new developments for TROPICS in the JEDI-GEOS atmospheric data assimilation system are in progress, and preliminary results from cycled JEDI-GEOS data assimilation experiments are presented.

Min-Jeong Kim↗
NASA NTRS1923

"Schneebeli" and "Birger" silencers

The Schneebeli silencer is made entirely of soft sheet steel. It is connected with the engine for receiving the exhaust gases, and consists of two parts: the silencer proper and a conical exit. The Birger silencer, made by the Ad Astra Company in Switzerland, is based on the principle of rapid cooling of the exhaust gases.

MUFFLERS↗
NASA NTRS2015

Automated Cooperative Trajectories for a More Efficient and Responsive Air Transportation System

The NASA Automated Cooperative Trajectories project is developing a prototype avionics system that enables multi-vehicle cooperative control by integrating 1090 MHz ES ADS-B digital communications with onboard autopilot systems. This cooperative control capability will enable meta-aircraft operations for enhanced airspace utilization, as well as improved vehicle efficiency through wake surfing. This briefing describes the objectives and approach to a flight evaluation of this system planned for 2016.

Wake Surfing↗
NASA NTRSDate not supplied

Formal Analysis of the Compact Position Reporting Algorithm

This presentation documents the formal analysis of the compact position reporting (CPR) algorithm. CPR is a fundamental part of Automatic Dependent Surveillance - Broadcast (ADS-B), which is a global protocol for aircraft communication. The formal analysis found and corrected issues with the algorithm, proposed simplifications, and created a formally verified reference implementation, all of which are incorporated in the governing standards document.

Formal Methods↗
NASA NTRSDate not supplied

Advancing Electric Propulsion Aircraft Evaluation for Urban Air Mobility: Insights from NASA Ames

This presentation delves into a recent evaluation conducted at NASA-Ames on the Vertical Motion Simulator, focusing on the handling qualities of Distributed Electric Propulsion VTOL (eVTOL) aircraft, specifically tailored for Urban Air Mobility (UAM) applications. The presentation will focus on the recent effort to adapt and refine use of the Aeronautical Design Standard -33 (ADS-33) rotorcraft handling qualities developed by the U.S. Army and NASA to meet the diverse needs of civilian (eVTOL) concept evaluation. A brief discussion of the author’s personal test pilot insights in the early development of military Fly-By-Wire evaluation methods will also be provided. The emergence of innovative eVTOL designs with unique lift capabilities and flight control systems, present both opportunities and challenges, particularly in ensuring safety amidst technological complexity. To navigate these challenges, our investigation examined evaluation criteria designed to accommodate the varied configurations and advanced automation systems inherent in modern eVTOL aircraft. By establishing a standardized approach to evaluation, our research not only fosters innovation but also upholds safety standards in the dynamic landscape of urban air mobility. Through this endeavor, we help to facilitate the seamless integration of novel aircraft capabilities into new operations, contributing to a new era of safe and efficient aerial transportation.

eVTOL↗
NASA NTRS2024

Interoperability Review of Wireless Hardware in Emergency Air-To-Ground Mesh Networks

The air-to-ground network has the goal of advancing the capabilities and versatility of communications in wildfire and emergency scenarios. To ensure compatibility and keep costs low for emergency services interoperability between hardware is an important step in development of the mesh network. For the purposes of this evaluation interoperability of hardware is described as the ability for hardware from different vendors to be interchangeable within the network with minimal initial setup and no functional change from an operator’s viewpoint. This report will also focus on the features and limitations in addition to the interoperability of the major wireless hardware systems core to the function of the proposed air-to-ground mesh network. The hardware systems involved are ADS-B receivers, 3GPP Sidelink devices, mesh radios, and commercial ISM band radios. For early testing the mesh radios will be in the 900 MHz frequency range using the DoodleLabs RM-1700-22W Mesh Rider Radio as the lab reference. This report will consider and categorize desirability of features, the amount of setup required before deployment, the effect on performance characteristics from nominal and maximum load, and the ability to interface with other vender devices.

Mesh Radio↗
NASA NTRS1962

Some Applications of Detailed Wind Profile Data to Launch Vehicle Response Problems

The response of a launch vehicle to a number of detailed wind profiles has been determined. The wind profiles were measured by two techniques which are briefly described. One of these techniques uses an angle-of-attack sensor in conjunction with guidance data to measure the wind profile traversed by some particular launch vehicle. The other wind-measuring technique is a photographic triangulation method, whereby two cameras take simultaneous pictures of a vertical trail of smoke left by a launch vehicle or sounding rocket. The response of a vehicle flying these detailed profiles is compared with the response of the same vehicle flying through balloon-measured profiles. The response to the detailed wind profiles, relative t o the balloon-measured profiles, is characterized by the large excitation of the rigid pitch and elastic bending modes. This is found to cause higher loads on the launch vehicle structure. Established design criteria which utilize balloon measured wind profiles have arbitrarily accounted-for this increased load by adding a load due to some type of discrete gust.

LAUNCH VEHICLE↗
NASA NTRS1962

Self erecting manned space laboratory

The Langley Research Center of the National Aeronautics and Space Administration is currently involved in several research programs on manned orbital space stations. This research effort is focused on seeking out and solving the problems which lie in the way of the eventual development of such a vehicle for use as a space laboratory. In conducting this research activity, it has been necessary to consider many vehicle configuration designs and operational concepts in some detail. The Langley Research Center space station study program has been in progress for several years, and it was consisted of both in-house activities and contracted efforts with industry. In the early Langley Research Center studies, only zero-gravity configurations were considered; however, the requirement for artificial gravity simulation for experimental purposes was soon added. Human factors considerations indicated that large diameter vehicles with slow rotational velocities generally permit the most comfortable living conditions for the crew. The configurations which were studied were required to be compatible with the planned launch vehicles such as Saturn, and manned spacecraft such as Mercury.

Manned Orbital Space Station (MOSS)↗
NASA NTRS1962

PROGRESS REPORT ON THE DEVELOPMENT OF PROTECTED CONSTRUCTION FOR HYPERSONIC VEHICLES

The structural problems of re-entry occasioned by aerodynamic heating are now generally well known. Typical of these heating problems is the environment experienced by the manned lifting vehicle re-entering from a low altitude orbit. Figure 1 shows typical curves of lower surface temperature as a function of time for a wing loading of 25 Ibs/ft(exp 2), and a lift-drag ratio of 2.5. The two curves cover a practical range of re-entry angles, the lower curve representing an ideal re-entry path with zero dive angle and the upper curve assuming a 2° error in the re-entry angle. Maneuvers for course change or correction are also added to the upper curve. Significant factors from this curve are the relationship between the maximum temperatures and the capabilities of available metallic material, and also the long re-entry time and its effect on total heat load. If a very shallow re-entry is made, maximum lower surface temperatures reach 2000°F which is just within the range of conventional superalloys. To accommodate practical re-entry angles and maneuvers, however, the temperature capability must reach about 2500°F which requires refractory metals. The re-entry time may be as high as 100 minutes, which gives total heat loads of approximately 40,000 BTU/ft(exp 2). A heat load of this magnitude, with equilibrium surface temperatures of the values shown, suggests that a lighter airframe can be constructed by dissipating the heat by radiation from the surface, rather than by absorbing it with a heat sink, or surface cooling, or ablation. Air Force programs to provide airframes for this type of environment have involved the parallel development of a number of different structural concepts. The development to be discussed here was carried out by Bell Aerosystems Company for the Fabrication Branch, Manufacturing Technology Laboratory, Directorate of Materials and Processes, Aeronautical Systems' Division, Wright-Patterson Air Force Base, Ohio, under Contract AF33(600)-40100 (Double-Wall Cooled Structure).

Reentry vehicle↗
NASA NTRS1962

DESIGN AND TEST EXPERIENCES WITH INSTABILITY OF MAJOR AIRFRAME COMPONENTS

Two test incidents involving instability of large scale commercially built structures are described. Two classes of structure are discussed; the first, a fuselage with skin designed to buckle at low stress, and the second, a wing whose surface remains unbuckled to failure. The structure in the region of failure is defined and the failures described and illustrated. Insofar as possible, the stresses in the critical area at the time of failure are reported and compared to strength determined by analysis. Both fuselage and wing surfaces were observed to fail in the mode of a medium range column when adequate support was provided by ribs and frames. Initial failure in the wing example was premature due to a design deficiency in rib strength. A clear illustration is given of the effect of rib stiffness on wing surface stability. Adding stiffness to wing ribs increases the limit of surface stability to the theoretical flat panel value.

SHELL STABILITY↗