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The inability of available situation displays to provide a decelerating instrument approach capability to a hover led to a flight research program in which control-command information was displayed for three degrees of freedom. The test aircraft (NASA's CH-46C in-flight simulator) was stabilized with high-gain attitude command system. Using this system, the pilot was able to decelerate the aircraft to a hover while simultaneously following a 6 deg glidepath. Although these tests demonstrate the potential of this concept, a number of factors, including adequate integration of command and situation information, were identified as affecting pilot acceptance of the system.

NASA's plans to implement the Vision for Space Exploration include extensive human-robot cooperation across an enterprise spanning multiple missions, systems, and decades. To make this practical, strong enterprise-level interface standards (data, power, communication, interaction, autonomy, and physical) will be required early in the systems and technology development cycle. Such standards should affect both the engineer and operator roles that humans adopt in their interactions with robots. For the engineer role, standards will result in reduced development lead-times, lower cost, and greater efficiency in deploying such systems. For the operator role, standards will result in common autonomy and interaction modes that reduce operator training, minimize workload, and apply to many different robotic platforms. Reduced quantities of spare hardware could also be a benefit of standardization.

The Plant Research Unit (PRU) is currently under development by the Space Station Biological Research Project (SSBRP) team at NASA Ames Research Center (ARC) with a scheduled launch in 2001. The goal of the project is to provide a controlled environment that can support seed-to-seed and other plant experiments for up to 90 days. This paper describes testing conducted on the major PRU prototype subsystems. Preliminary test results indicate that the prototype subsystem hardware can meet most of the SSBRP science requirements within the Space Station mass, volume, power and heat rejection constraints.

Firefighters want to go to work, do their job well, and go home alive and uninjured. For their most important job, saving lives, firefighters want protective equipment that will allow more extended and effective time at fire scenes in order to perform victim search and rescue. A team, including engineers at NASA JSC and firefighters from Houston, has developed a list of problem areas for which NASA technology and know-how can recommend improvements for firefighter suits and gear. Prototypes for solutions have been developed and are being evaluated. This effort will spin back to NASA as improvements for lunar and planetary suits.

The Orion Crew Exploration Vehicle (CEV) requires a smoke detector for the detection of particulate smoke products as part of the Fire Detection and Suppression (FDS) system. The smoke detector described in this paper is an adaptation of a mature commercial aircraft design for manned spaceflight. Changes made to the original design include upgrading the materials and electronics to space-qualified components, and modifying the mechanical design to withstand launch and landing loads. The results of laboratory characterization of the response of the new design to test particles are presented.

Unlike a terrestrial electric utility which can purchase power from a neighboring utility, the Space Station Freedom (SSF) has strictly limited energy resources; as a result, source control, system monitoring, system protection and load management are essential to the safe and efficient operation of the SSF Electric Power System (EPS). These functions are being evaluated in the DC Power Management and Distribution (PMAD) Testbed which NASA LeRC has developed at the Power System Facility (PSF) located in Cleveland, Ohio. The testbed is an ideal platform to develop, integrate, and verify power system monitoring and control algorithms. State Estimation (SE) is a monitoring tool used extensively in terrestrial electric utilities to ensure safe power system operation.

Spacecraft hardware trade studies compare options primarily on mass while considering impacts to cost, risk, and schedule. Historically, other factors have been considered in these studies, such as reliability, technology readiness level (TRL), volume and crew time. In most cases, past trades compared two or more technologies across functional and TRL boundaries, which is an uneven comparison of the technologies. For example, low TRL technologies with low mass were traded directly against flight-proven hardware without consideration for requirements and the derived architecture. To provide for even comparisons of spacecraft hardware, trades need to consider functionality, mission constraints, integer vs. real number of flight hardware units, and mass growth allowances by TRL.

Due to numerous engine power-loss events associated with high-altitude convective weather, ice accretion within an engine due to ice-crystal ingestion is being investigated. The National Aeronautics and Space Administration (NASA) and the National Research Council (NRC) of Canada are starting to examine the physical mechanisms of ice accretion on surfaces exposed to ice-crystal and mixed-phase conditions. In November 2010, two weeks of testing occurred at the NRC Research Altitude Facility utilizing a single wedge-type airfoil designed to facilitate fundamental studies while retaining critical features of a compressor stator blade or guide vane. The airfoil was placed in the NRC cascade wind tunnel for both aerodynamic and icing tests. Aerodynamic testing showed excellent agreement compared with CFD data on the icing pressure surface and allowed calculation of heat transfer coefficients at various airfoil locations.

During the past two decades the occurrence of ice accretion within commercial high bypass aircraft turbine engines under certain operating conditions has been reported. Numerous engine anomalies have taken place at high altitudes that were attributed to ice crystal ingestion such as degraded engine performance, engine roll back, compressor surge and stall, and even flameout of the combustor. As ice crystals are ingested into the engine and low pressure compression system, the air temperature increases and a portion of the ice melts allowing the ice-water mixture to stick to the metal surfaces of the engine core. The focus of this paper is on estimating the effects of ice accretion on the low pressure compressor, and quantifying its effects on the engine system throughout a notional flight trajectory. In this paper it was necessary to initially assume a temperature range in which engine icing would occur.

The value of “ultracapacitors” (also referred to as “supercapacitors” or “electric double layer capacitors” in some literature) as an augmentation device when placed in parallel with “electrochemical” energy storage (i.e. battery) is presented in this paper. Since ultracapacitors possess unique attributes due to their higher value of energy storage density (or Joules/WattHrs per mass) compared to conventional capacitors while maintaining the peak power providing capability (to some degree) typical of conventional capacitors they may provide a near term solution in applications demanding longer battery operating life when placed in parallel. Such demands may be pronounced by the onset of More-Electric-Aircraft peak loads and “cold-crank” Auxiliary Power Unit (APU) electric-starting in demanding cold temperature environments.

In this investigation, analytical models were developed to predict the performance characteristics of axisymmetric single jet ejector. The ejector is divided into four parts, jet, mixer, nozzle, and diffuser. Basic flow equations were combined to calculate end to end flow characteristics for each of the four ejector components. Different jets and mixer combination were tested using three jet and three mixers. Characteristics curves have been drawn to predict flow characteristics of the ejector. Different configuration of jet and mixer incorporated different loss coefficient. Hence to get correct flow characteristics of the ejector right loss coefficient should be used.

Advancements in electrical, mechanical, and structural design onboard modern more electric aircraft have added significant stress to the thermal management systems (TMS). A thermal management system level analysis tool has been created in MATLAB/Simulink to facilitate rapid system analysis and optimization to meet the growing demands of modern aircraft. It is anticipated that the tracking of thermal energy through numerical integration will lead to more accurate predictions of worst case TMS sizing conditions. In addition, the non-proprietary nature of the tool affords users the ability to modify component models and integrate advanced conceptual designs that can be evaluated over multiple missions to determine the impact at a system level.

Under a NASA-sponsored technology development project, a multi-disciplinary team consisting of industry, academia, and government organizations lead by Hamilton Sundstrand is developing an amine-based humidity and CO2 removal process and prototype equipment for Vision for Space Exploration (VSE) applications. Originally this project sought to research enhanced amine formulations and incorporate a trace contaminant control capability into the sorbent. In October 2005, NASA re-directed the project team to accelerate the delivery of hardware by approximately one year and emphasize deployment on board the Crew Exploration Vehicle (CEV) as the near-term developmental goal. Preliminary performance requirements were defined based on nominal and off-nominal conditions and the design effort was initiated using the baseline amine sorbent, SA9T.

NASA and the U.S. Army have designed, developed, and tested a Computer Aiding for Low-Altitude Helicopter Flight guidance system. This system provides guidance to the pilot for near-terrain covert helicopter operations. The guidance is presented to the pilot through symbology on a helmet mounted display. This system has demonstrated the feasibility of a pilot-centered concept of terrain flight guidance that preserves pilot flexibility and authority. The system was developed using extensive piloted simulation and then implemented in a UH-60 Blackhawk helicopter for flight development and evaluation. A close correlation between simulation and actual flight was found; however, in flight overall pilot workload increased and performance decreased. This paper presents a description of the basic system design, simulation, and flight evaluations.

The Carbon Dioxide Removal Assembly (CDRA) on board the International Space Station (ISS) has experienced periodic check valve and selector valve failures as a result of a gradual build-up of contamination from particles that have breeched the adsorbent bed seals. The current software that controls CDRA has limitations that make troubleshooting the unit difficult in these situations, in large part due to the fact that valve position telemetry is only available during certain times. There are also situations where it is required to perform operations manually that would benefit from added code logic and commands to facilitate these operations. The software has been reviewed for possible upgrades and changes that will allow engineers to better troubleshoot the unit in the event of various failures and also allow for better operability in degraded states.

Significant system efficiency gains can be achieved in high-performance aircraft via a unitized structure that reduces parts count. For instance, reduced parts count leads to substantial engineering logistic cost savings through higher levels of subsystem and mounting hardware integration. It also creates performance benefits by eliminating structural connections. Residual stress management, however, remains a major obstacle to capturing full benefits and broadening the application of unitized structure solutions. This paper describes how Alcoa and others are developing tools to overcome limitations in current testing, evaluation, and design practices attributed to residual stress effects. The authors present recent advancements in fracture toughness and fatigue crack growth characterization, along with a new, integrated approach for improved accounting of residual stress effects during fracture critical component design, manufacturing planning, and life management.

A dynamic simulation of the ISS CDRA hardware was created using the Aspen Custom Modeler software platform. The dynamic model calculates the material and energy balances that describe the system properties. The model was calibrated by comparison to test data results from a flight-like CDRA at NASA Marshall Space Flight Center. While other FORTRAN models of the CDRA already exist, developing an ACM simulation is the first step towards creating a generic tool to simulate larger collections of life support hardware. The ACM tool should make it possible to be very flexible when rearranging these models to simulate possible configurations of the life support subsystems that could be used in the future, especially for advanced life support applications.

The Lunar Electric Rover (LER), which was formerly called the Small Pressurized Rover (SPR), is currently being carried as an integral part of the lunar surface architectures that are under consideration in the Constellation Program. One element of the LER is the suit port, which is the means by which crew members perform Extravehicular Activities (EVAs). Two suit port deliverables were produced in fiscal year 2008: a 1-g suit port concept evaluator for functional integrated testing with the LER 1-g concept vehicle and a functional and pressurizable Engineering Unit (EU). This paper focuses on the 1-g suit port concept evaluator test results from the Desert Research and Technology Studies (D-RATS) October 2008 testing at Black Point Lava Flow (BPLF), Arizona. The 1-g suit port concept evaluator was integrated with the 1-g LER cabin and chassis concepts.