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An integration study was performed coupling an SP-100 reactor with either a Brayton or Stirling power conversion subsystem. A power level of 100 kWe was selected for the study. The power system was to be compatible with both the lunar and Mars surface environment and require no site preparation. In addition, the reactor was to have integral shielding and be completely self-contained, including its own auxiliary power for start-up. Initial reliability studies were performed to determine power conversion redundancy and engine module size. Previous studies were used to select the power conversion optimum operating conditions (ratio of hot-side temperature to cold-side temperature). Results of the study indicated that either the Brayton or Stirling power conversion subsystems could be integrated with the SP-100 reactor for either a lunar or Mars surface power application.

Sundstrand has been investigating 270-Vdc/hybrid 115-Vac electrical power generating systems (EPGS) technology in preparation for meeting the electrical power generating system (EPGS) requirements for future aircraft (1). Systems such as the one being investigated are likely to be suitable for the More-Electric Aircraft (MEA) concepts presently under industry and military study. The present Sundstrand single-channel testbed is being further expanded to better understand the electrical system performance characteristics and power quality requirements of an MEA in which traditional mechanical subsystems are replaced by those of a “more-electric” nature. This paper presents the most recent Sundstrand 270-Vdc system transient performance data, and describes the modifications being made to the 270-Vdc/hybrid 115-Vac testbed.

Different Airbus Helicopters main rotor blade profiles were tested in different icing wind tunnels and for different icing conditions. One of the objectives of the accretion tests was to validate the use of 2D icing scaling laws established for fixed wing aircraft on helicopter blade profiles. Therefore, ice shapes resulting from tests with the same icing similarity parameters are compared to each other allowing the assessment of icing scaling laws for helicopter applications. This paper presents the icing scaling laws used at Airbus Helicopters on blade profiles, the different test set ups and test models and it presents the comparison of the ice shapes collected during the icing wind tunnel test campaigns.

A helicopter blade profile was tested in the DGA Aero-engine Testing's icing altitude test facility S1 in Saclay, France during the winter of 2013/2014. The airfoil was a helicopter main rotor OA312 blade profile made out of composite material and with a metallic erosion shield. Dry air and ice accretion tests have been performed in order to assess the iced airfoil's aerodynamic behaviour. Several icing conditions were tested up through Mach numbers around 0.6. This paper presents the test setup, the test model and some of the test results. The test results presented in this paper include the ice shapes generated as well as dry air and iced airfoil lift and drag curves (polars) which were obtained with the real ice shapes on the airfoil.

The term “3 inch ice shapes” has assumed numerous definitions throughout the years. At times it has been used to generally characterize large glaze ice accretions on the major aerodynamic surfaces (wing, horizontal stabilizer, vertical stabilizer) for evaluating aerodynamic performance and handling qualities after a prolonged icing encounter. It has also been used as a more direct criterion while determining or enforcing sectional ice shape characteristics such as the maximum pinnacle height. It is the authors’ observation that over the years, the interpretation and application of this term has evolved and is now broadly misunderstood. Compounding the situation is, at present, a seemingly contradictory set of guidance among (and even within) the various international regulatory agencies resulting in an ambiguous set of expectations for design and certification specialists.

A 3D computer model named AIPAC (Aircraft Ice Protection Analysis Code) suitable for thermal ice protection system parametric studies has been developed. It was derived from HASPAC, which is a 2D anti-icing model developed at Wichita State University in 2010. AIPAC is based on the finite volumes method and, similarly to HASPAC, combines a commercial Navier-Stokes flow solver with a Messinger model based thermodynamic analysis that applies internal and external flow heat transfer coefficients, pressure distribution, wall shear stress and water catch to compute wing leading edge skin temperatures, thin water flow distribution, and the location, extent and rate of icing. In addition, AIPAC was built using a transient formulation for the airfoil wall and with the capability of extruding a 3D surface grid into a volumetric grid so that a layer of ice can be added to the computational domain.

The in-flight ice accretion simulations are typically performed using a quasi-steady formulation through a multi-step approach. As the ice grows, the geometry changes, and an adaptation of the fluid volume mesh used by the airflow and droplet-trajectory solver is required. Re-meshing or mesh deformation are generally employed to do that. The geometries formed are often complex ice shapes increasing the difficulty of the re-meshing process, especially in three-dimensional simulations. Consequently, difficulties are encountered when trying to automate the process. Contrary to the usual body-fitted mesh approach, the use of immersed boundary methods (IBMs) allows solving, or greatly reducing, this problem by removing the mesh update, facilitating the global automation of the simulation. In the following paper, an approach to perform the airflow and droplet trajectory calculations for three-dimensional simulations is presented. This framework utilizes only immersed boundary methods.

Vertical-Junction-Field-Effect-Transistors (VJFETs) are currently the most mature SiC devices for high power/temperature switching. High-voltage VJFETs are typically designed normally-on to ensure voltage control operation at high current-gain. However, to exploit the high voltage/temperature capabilities of VJFETs in a normally-off high-current voltage-controlled switch, high-voltage normally-on and low-voltage normally-off VJFETs were connected in the cascode configuration. In this paper, we review the high temperature DC characteristics of VJFETs and 1200 V normally-off cascode switches. The measured parameter shifts in the 25°C to 300°C temperature range are in excellent agreement with theory, confirming fabrication of robust SiC VJFETs and cascode switches.

The 757/767 Flight Management System provides the initial operational implementation of an integrated guidance, control and display equipments based upon digital technology for commercial transport airplanes. The applied equipments are based upon the new ARINC 700 series characteristics developed by the Industry over the past five years. These characteristics were developed on the basis of limited operational experience with selected elements of the system and upon R&D efforts within the Industry. The System features automatic/manual flight profiles for optimum economics, all weather landing including rollout guidance, electronic primary flight instruments based on color (shadow mask) CRTs, inertial attitude/velocity reference based upon laser gyros, improved caution/warning and other improved performance/functional features. The system also provides significant improvements in line and shop maintenance features.

This paper describes the wing and engine ice protection system, used on all 777 aircraft. The 777 ice protection system is unique in two ways: it has an advanced control system which minimizes aircraft power consumption. In addition, the system was procured by the prime contractor, Boeing, as a fully integrated subsystem from a single supplier.

In the course of CRYOSYSTEM phase B (development phase) financed by the European Space Agency, AIR LIQUIDE (France) and Astrium Space Infrastructure (Germany) have developed an optimized design of a −183°C freezer to be used on board the International Space Station for the freezing and storage of biological samples. The CRYOSYSTEM facility consists of the following main elements: - the CRYORACK, an outfitted standard payload rack (ISPR) accommodating up to three identical Vial Freezers - the Vial Freezer, a dewar vessel capable of fast and ultra-rapid freezing, and storing up to approximately 900 vials below −183°C; the dewar is cooled by a Stirling machine producing > 6 W at 90 K. The Vial Freezer is operational while accommodated in the CRYORACK or attached to the Life Science Glovebox (LSG). One CRYORACK will remain permanently on-orbit for several years while four Vial Freezers and two additional CRYORACKs support the cyclic upload/download of samples.

Some of the methods used for experimental detection and examination of wake vortices are presented. The aim of the article is to provide the reader a brief overview of the available methods. The material is divided into two major sections, one dealing with methods used primarily in the laboratory, and the second part devoted to those used in field operations. Over one hundred articles are cited and briefly discussed.

Europe has embarked on a new programme of space exploration involving the development of rover, lander and probe missions to visit planets, moons and near Earth objects (NEOs) throughout the Solar System. Rovers and landers will require testing under simulated planetary, and NEO conditions to ensure their ability to land on and traverse the alien surfaces. ESA has begun work on a building project that will provide an enclosed and controlled environment for testing rover and lander functions such as landing, mobility, navigation and soil sampling. The facility will first support the European ExoMars mission due for launch in 2013. This mission will deliver a robotic rover to the Martian surface. This paper, the first of several on the project, gives an overview of its design configuration and construction phasing. Future papers will cover its applications and operations.

In this work, a newly developed iced-aircraft modeling tool is applied to wings, engine inlets, and helicopter rotors. The tool is based on a multiscale-physics, unstructured finite-volume CFD approach and is applicable to general purpose aircraft icing applications. The present approach combines an Eulerian-based droplet-trajectory solver that is loosely coupled, in a time-accurate manner, to a surface-film and ice-evolution model. The goal of the model is to improve the fidelity of ice accretion modeling on dynamic geometries and for three-dimensional ice shapes typical of helicopter rotors. The numerical formulation is discussed and presented alongside 2D and 3D static validation cases, and dynamic helicopter rotors. The present results display good validation for predicting ice shape on a variety of geometries, and a strong initial capability of modeling ice forming on helicopters in forward flight.

This work presents the implementation and validation efforts of a 3D ice accretion solver for aeronautical applications, MESS3D, based on the advanced Messinger model. The solver is designed to deal with both liquid phase and ice crystal cloud conditions. In order to extend the Messinger model to 3D applications, an algorithm for the water run-back distribution on the surface was implemented, in place of an air flow stagnation line search algorithm, which is straightforward in 2D applications, but more complicated in 3D. The developed algorithm aims to distribute the run-back water in directions determined by air pressure gradients or shear forces. The data structure chosen for MESS3D allows high flexibility since it can manage the necessary input solutions on surface grids coming from both structured and unstructured solvers, regardless the number of edges per surface cells.

The objectives are to compare different psychological methods used to assess the evolution of the interrelations inside the crew and the relationships between the crew and the outside in a sixty days isolation/confinement's simulation. After presenting each method, results are compared. The discussion try to point out if these methods are equivalent or if they are complementary. The specificity of each method is shown and conclusions try to associate some methods with specific scientific goals.

Two turbulence models have been studied to determine which of the models should be used in further Computational Fluid Dynamics (CFD) research. A zero-equation turbulence model, Baldwin-Lomax (B-L), is easy to use, requires no history of the flow, and requires little in the way of additional computations or additional computer memory space [1]. A two-equation k-ε model, Yang-Shih (Y-S), is more difficult to implement, does require flow history, and requires many more computations and much more computer space; however, it is potentially more accurate than the B-L model [2]. Using both Navier-Stokes (NS) and Parabolized Navier-Stokes (PNS) solvers, the two models and their codes were validated against the testbed of the Wright Laboratory (WL) Mach 12 wind tunnel nozzle.

There are many different numerical approaches to ice accretion simulation. Little comparison has been made between those approaches to identify the best tool for a given application. This paper presents a comparison exercise between 2D codes (CANICE-BA and LEWICE) and 3D codes (CANICE3D-BA, LEWICE3D and FENSAP-ICE). It also compares the 3D first generation code (panel method with Lagrangian droplet trajectory tracking) CANICE3D-BA to the 3D second generation code (Navier-Stokes with Eulerian droplet tracking) FENSAP-ICE. The paper includes a description of the different methodologies. The first comparison exercise is done using three 2D cases for which experimental ice shapes are available. The second exercise addresses a water collection efficiency over an isolated tail for which experimental data is available. Finally, an ice accretion comparison is presented in a DLR4 wing-body configuration.

Multiphase CFD simulations of air and water play a critical role in aircraft icing analysis. Specifically for air data sensors mounted near the front of an aircraft, simulations that predict the concentration of water surrounding an aircraft fuselage are necessary for understanding their performance in icing conditions. Those simulations can aid in sensor design and placement, and are central for defining critical conditions to test during icing qualification campaigns. There are several methods available in CFD that solve a multiphase flow field. Two of the most common methods used are Lagrangian and Eulerian. While these methods are similar, important differences can be viewed in the results, specifically in how the water shadow zones are predicted. This paper compares a Lagrangian and Eulerian CFD method for solving a multiphase flow field, and assesses their performance for use for analyzing installation locations and critical icing conditions of air data probes.