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Operational issues encountered by Apollo astronauts relating to lunar dust were catalogued, including material abrasion that resulted in scratches and wear on spacesuit components, ultimately impacting visibility, joint mobility and pressure retention. Standard methods are being developed to measure abrasive wear on candidate construction materials to be used for spacesuits, spacecraft, and robotics. Calibration tests were conducted using a standard diamond stylus scratch tip on the common spacecraft structure aluminum, Al 6061-T6. Custom tips were fabricated from terrestrial counterparts of lunar minerals for scratching Al 6061-T6 and comparing to standard diamond scratches. Considerations are offered for how to apply standards when selecting materials and developing dust mitigation strategies for lunar architecture elements.

The National Aeronautics and Space Administration (NASA) conducted a full scale ice crystal icing turbofan engine test using an obsolete Allied Signal ALF502-R5 engine in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The test article used was the exact engine that experienced a loss of power event after the ingestion of ice crystals while operating at high altitude during a 1997 Honeywell flight test campaign investigating the turbofan engine ice crystal icing phenomena. The test plan included test points conducted at the known flight test campaign field event pressure altitude and at various pressure altitudes ranging from low to high throughout the engine operating envelope. The test article experienced a loss of power event at each of the altitudes tested.

The On-Board Diagnostics II (OBD-II) port began as a means of extracting diagnostic information and supporting the right to repair. Self-driving vehicles and cellular dongles plugged into the OBD-II port were not anticipated. Researchers have shown that the cellular modem on an OBD-II dongle may be hacked, allowing the attacker to tamper with the vehicle brakes. ADAS, self-driving features and other vehicle functions may be vulnerable as well. The industry must balance the interests of multiple stakeholders including Original Equipment Manufacturers (OEMs) who are required to provide OBD function, repair shops which have a legitimate need to access the OBD functions, dongle providers and drivers. OEMs need the ability to protect drivers and manage liability by limiting how a device or software application may modify the operation of a vehicle.

Atmospheric Neutron Single Event Effects (SEE) are widely known to cause failures in all electronic hardware, and cause proportionately more failures in avionics equipment due to the use altitude. In digital systems it is easy to show how SEE can contribute several orders of magnitude more faults than random (hard) failures. Unfortunately, current avionics Safety assessment methods do not require consideration of faults from SEE. AVSI SEE Task Group (Aerospace Vehicle Systems Institute Committee #72, on Mitigating Radiation Effects in Avionics) is currently coordinating development of an atmospheric Neutron Single Event Effects (SEE) Analysis method. This analysis method is a work in progress, in close collaboration with SAE S-18 and WG-63 Committees (Airplane Safety Assessment Committee). The intent is to include this method as part of current revisions to ARP4761 (Guidelines and Methods for Conducting the Safety Assessment Process on Civil Airborne Systems and Equipment).

Natural atmospheric radiation effects have been recognized in recent years as key safety and reliability concerns for avionics systems. Atmospheric radiation may cause Single Event Effects (SEE) in electronics. The resulting Single Event Effects can cause various fault conditions, including hazardous misleading information and system effects in avionics equipment. As technology trends continue to achieve higher densities and lower voltages, semiconductor devices are becoming more susceptible to atmospheric radiation effects. To ensure a system meets all its safety and reliability requirements, SEE induced upsets and potential system failures need to be considered. The purpose of this paper is to describe a process to incorporate the SEE analysis into the development like-cycle. Background on the atmospheric radiation phenomenon and the resulting single event effects, including single event upset (SEU) and latch up conditions is provided.

In applications ranging from design of customized vehicle interiors to virtual testing of biomedical devices, the processes of modeling, design and analysis involve the simultaneous treatment of artifacts (i.e., parts designed by humans) and anatomical structures. An inherent conflict arises because the geometric descriptions are completely different. Artifact descriptions are typically the output of computer-aided design (CAD) software and consist of a collection of parametric patches that comprise the boundary of the artifact. In stark contrast, the native description of an anatomical structure typically consists of an image stack obtained using a volumetric scanning technology such as computed tomography (CT) or magnetic resonance imaging (MRI). Current practice for simultaneously dealing with both categories of entities involves working primarily in the world of CAD.

In order to reduce painful and injurious shear stresses for lower limb amputees, prosthetic ankle joints need to provide torsional control in the transverse plane. This paper attempts to characterize biological ankle function in the transverse plane with simple mechanical elements to assist in the design of a biomimetic prosthetic ankle joint. Motion capture data was collected from ten subjects walking in a straight trajectory to model four states of stance phase. Passive elements were chosen to model the ankle in each state. The ankle was observed to act as a quadratic torsional spring in State 1 and as linear torsional springs in States 2, 3 and 4. The results of this study should assist with the mechanical design and control of a biomimetic torsional prosthesis by suggesting a finite state control system and by providing the stiffness coefficients to be controlled for straight walking.

The NASA-wide CEV Smart Buyer Team (SBT) was assembled in January 2006 and was tasked with the development of a NASA in-house design for the CEV Crew Module (CM), Service Module (SM), and Launch Abort System (LAS). This effort drew upon over 250 engineers from all of the 10 NASA Centers. In 6 weeks, this in-house design was developed. The Thermal Systems Team was responsible for the definition of the active and passive design architecture. The SBT effort for Thermal Systems can be best characterized as a design architecting activity. Proof-of-concepts were assessed through system-level trade studies and analyses using simplified modeling. This nimble design approach permitted definition of a point design and assessing its design robustness in a timely fashion. This paper will describe the architecting process and present trade studies and proposed thermal designs

Innovative multi-environment multimode thermal management architecture has been described that is capable of meeting widely varying thermal control requirements of various exploration mission scenarios currently under consideration. The proposed system is capable of operating in a single-phase or two-phase mode rejecting heat to the colder environment, operating in a two-phase mode with heat pump for rejecting heat to a warm environment, as well as using evaporative phase-change cooling for the mission phases where the radiator is incapable of rejecting the required heat. A single fluid loop can be used internal and external to the spacecraft for the acquisition, transport and rejection of heat by the selection of a working fluid that meets NASA safety requirements. Such a system may not be optimal for each individual mode of operation but its ability to function in multiple modes may permit global optimization of the thermal control system.

This paper presents a review of our current experimental and theoretical understanding of icing feathers and the role that they play in the formation of ice accretions. It covers the following areas: a short review of past research work related to icing feathers; a discussion of the physical characteristics and terminology used in describing icing feathers; the presence of feathers on ice accretions formed in unswept airfoils, especially at SLD conditions; the role that icing feathers play in the formation of ice accretion shapes on swept wings; the formation of icing feathers from roughness elements; theoretical considerations regarding feather formation, feather interaction to form complex icing structures, the role of film dynamics in the formation of roughness elements and the formation of feathers. Hypotheses related to feather formation and feather growth are discussed.

An experiment was conducted in the Icing Research Tunnel (IRT) at NASA Glenn Research Center to study the effect of sweep angle and temperature on the formation of ice accretions on a NACA 0012 swept wing at SLD conditions. From a baseline Appendix-C condition with a MVD of 20m the drop size was changed to 110 and 200m for the SLD cases. Casting data, ice shape tracings, time-sequence and photographic data were obtained. Time-sequence photography was taken during each run to capture in real time the formation of the ice accretion. Measurements of the critical distance were obtained.

Map matching determines which road a vehicle is on based on inaccurate measured locations, such as GPS points. Simple algorithms, such as nearest road matching, fail often. We introduce a new algorithm that finds a sequence of road segments which simultaneously match the measured locations and which are traversable in the time intervals associated with the measurements. The time constraint, implemented with a hidden Markov model, greatly reduces the errors made by nearest road matching. We trained and tested the new algorithm on data taken from a large pool of real drivers.

This paper develops and tests algorithms for predicting the end-to-end route of a vehicle based on GPS observations of the vehicle's past trips. We show that a large portion a typical driver's trips are repeated. Our algorithms exploit this fact for prediction by matching the first part of a driver's current trip with one of the set of previously observed trips. Rather than predicting upcoming road segments, our focus is on making long term predictions of the route. We evaluate our algorithms using a large corpus of real world GPS driving data acquired from observing over 250 drivers for an average of 15.1 days per subject. Our results show how often and how accurately we can predict a driver's route as a function of the distance already driven.

This work deals with the modeling and analysis of a phasor-controlled Starter/Generator (S/G) electrical machine during starting either an aircraft Main Engine (ME) or Auxiliary Power Unit (APU). The model can be used to determine how much stator and exciter current is required to be supplied by a controlled power converter to the S/G to meet the start torque profile. In addition to modeling details and simulation results the paper presents a thorough analysis of the S/G machine, its environment and control.

This work deals with the modeling and analysis of both AC and DC bus voltage control in aerospace applications. The results of the analysis are presented along with system models, including a voltage-controlled current source (vccs) used as a DC Bus controller, a d,q-controlled, IGBT-based, SVPWM-switched, ac-to-dc active converter/rectifier (AR) used as a DC Bus controller, a 3-phase ac generator voltage regulator (VR) used as an AC Bus controller, a 3-phase uncontrolled ac generator followed by an SCR-controlled ac-to-dc converter, used as a DC Bus controller (single-controlled bus), and a 3-phase dynamically-controlled ac generator followed by an SCR-controlled ac-to-dc converter, used to provide both AC and DC Bus control (dual-controlled bus).

NASA and the FAA are funding the development of ground-based remote sensing systems specifically designed to detect and quantify the icing environment aloft. The goal of the NASA activity is to develop a relatively low cost stand-alone system that can provide practical icing information to the flight community. The goal of the FAA activity is to develop more advanced systems that can identify supercooled large drop (SLD) as well as general icing conditions and be integrated into the existing weather information infrastructure. Both activities utilize combinations of sensing technologies including radar, radiometry, and lidar, along with Internet-available external information such as numerical weather model output where it is found to be useful. In all cases the measured data of environment parameters will need to be converted into a measure of icing hazard before it will be of value to the flying community.

This paper discusses the problem of designing electric machines (EM) for advanced electric generators (AEG) used in aerospace more electric architecture (MEA) that would be applicable to aircraft, spacecraft, and military ground vehicles. The AEG's are analyzed using aspects of Six Sigma theory that relate to critical-to-quality (CTQ) subjects. Using this approach, weight, volume, reliability, efficiency, and cost (CTQs) are addressed to develop a balance among them, resulting in an optimized power generation system. The influence of the machine power conditioners and system considerations are also discussed. As a part of the machine evaluation process, speeds, bearings, complexities, rotor mechanical and thermal limitations, torque pulsations, currents, and power densities are also considered. A methodology for electric machine selection is demonstrated. Examples of high-speed, high-performance machine applications are shown.

To support NASA's goal to improve aviation safety, the Aircraft Icing Project of the Aviation Safety Program has developed a number of education and training aids for pilots and operators on the hazards of atmospheric icing. A review of aircraft incident and accident investigations has revealed that flight crews have not always understood the effects of ice contamination on their aircraft. To increase this awareness, NASA has partnered with regulatory agencies and pilot trade organizations to assure relevant and practical materials that are focused toward the intended pilot audience. A number of new instructional design approaches and media delivery methods have been introduced to increase the effectiveness of the training materials by enhancing the learning experience, expanding user interactivity and participation, and, hopefully, increasing learner retention rates.

This paper reviews work conducted in the UK aimed at developing validated methods to simulate ice accretion formed in super-cooled large droplet (SLD) icing conditions. To date, QinetiQ has completed one theoretical and three experimental programmes of work. Two further studies are currently in progress within UK universities. This paper provides results from the third test conducted by QinetiQ and NASA in the GKN Aerospace Composite Technologies Icing Research Wind Tunnel, Luton UK, to measure the mass loss through droplet splash during an SLD encounter. A description of the test procedures and the results obtained are provided. Future work on SLD methods development in progress in the UK is then briefly outlined.

This paper presents experimental methods for investigating large droplet impingement dynamics and for obtaining small and large water droplet impingement data. Droplet impingement visualization experiments conducted in the Goodrich Icing Wind Tunnel with a 21-in chord NACA 0012 airfoil demonstrated considerable droplet splashing during impingement. The tests were performed for speeds in the range 50 to 175 mph and with cloud median volumetric diameters in the range of 11 to 270 microns. Extensive large droplet impingement tests were conducted at the NASA Glenn Icing Research Tunnel (IRT). Impingement data were obtained for a range of airfoil sections including three 36-inch chord airfoils (MS(1)-0317, GLC-305, and NACA 652-415), a 57-inch chord Twin Otter horizontal tail section and 22.5-minute and 45-minute LEWICE glaze ice shapes for the Twin Otter tail section. Small droplet impingement tests were also conducted for selected test models.