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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.

This Specification defines general architectural philosophy and aircraft infrastructure for the proper use and interface of various cabin related IFE equipment. Compliance with ARINC Specification 808 allows each respective system to operate in concert when integrated with other relevant cabin equipment. ARINC Specification 808 defines standards for the aircraft 3rd Generation Cabin Network (3GCN), IFE Cabin Distribution System (CDS), wiring, connectors, power, identification codes, space envelopes, and mounting principles. Although some of these standards also apply to 3GCN wireless IFE systems, the overall 3GCN wireless IFE network specification is covered in ARINC Specification 820. The equipment itself is not a subject of this specification because it may be unique to the system manufacturer or marketplace-driven. Design guidelines are included for informational purposes as these guidelines impact the interfaces and installation of cabin equipment aboard the aircraft.

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 transition towards More Electric Aircraft (MEA) architectures has challenges relating to integration of power electronics with the starter generator system for on-engine application. To efficiently operate the power electronics in the hostile engine environment at high switching frequency and for better thermal management, use of silicon carbide (SiC) power devices for a bi-directional power converter is examined. In this paper, development of a 50 kVA bi-directional converter operating at an ambient temperature of about 2000C is presented. The design and operation of the converter with details of control algorithm implementation and cooling chamber design are also discussed.

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.

An electrical system composed of an integral main engine electrical starter/generator and a power management and distribution system for a high performance more electric airplane is presented. The paper emphasizes the fault tolerance and high reliability requirements and discusses the necessary architecture to satisfy them. A number of critical technical issues associated with the more electric airplane, such as severe EMI environment, regenerative power, and system integration and stability are discussed.

Abstract Identity-Anonymized CAN (IA-CAN) protocol is a secure CAN protocol, which provides the sender authentication by inserting a secret sequence of anonymous IDs (A-IDs) shared among the communication nodes. To prevent malicious attacks from the IA-CAN protocol, a secure and robust system error recovery mechanism is required. This article presents a central management method of IA-CAN, named the IA-CAN with a global A-ID, where a gateway plays a central role in the session initiation and system error recovery. Each ECU self-diagnoses the system errors, and (if an error happens) it automatically resynchronizes its A-ID generation by acquiring the recovery information from the gateway. We prototype both a hardware version of an IA-CAN controller and a system for the IA-CAN with a global A-ID using the controller to verify our concept.

In an effort to help future crewed spacecraft thermal control analysts understand the characteristics of one-and two-loop Active Thermal Control Systems (ATCS), a comparison was made between the one- and two-loop ATCS architectures officially proposed for the Crew Exploration Vehicle (CEV) in Design Analysis Cycle 1 (DAC1) and DAC2, respectively. This report provides a description of each design, along with mass and power estimates derived from their respective Master Equipment List (MEL) and Power Equipment List (PEL). Since some of the components were sized independent of loop architecture (ex. coldplates and heat exchangers), the mass and power for these components were based on the MEL and PEL of the most mature design (i.e. two-loop architecture). The mass and power of the two architectures are then compared and the ability of each design to meet CEV requirements is discussed.

The space exploration enterprise that will lead to human exploration on Mars requires pressure suit system capabilities and characteristics that change significantly over time and between different missions and mission phases. These capabilities must be provided within tight budget constraints and severely limited launch mass and volume, and at a pace that supports NASA's over-all exploration timeline. As a result, it has not been obvious whether the use of a single pressure suit system (like Apollo) or combinations of multiple pressure suit designs (like Shuttle) will offer the best balance among life cycle cost, risk, and performance. Because the answer to this question is pivotal for the effective development of pressure suit system technologies that will met NASA's needs, ILC and Hamilton Sundstrand engineers have collaborated in an independent study to identify and evaluate the alternatives.

The Orion Crew and Service Modules are often compared to the Apollo Command and Service Modules due to their similarity in basic mission objective: both were dedicated to getting a crew to lunar orbit and safely returning them to Earth. Both spacecraft rely on the environmental control, life support and active thermal control systems (ECLS/ATCS) for the basic functions of providing and maintaining a breathable atmosphere, supplying adequate amount of potable water and maintaining the crew and avionics equipment within certified thermal limits. This assessment will evaluate the driving requirements for both programs and highlight similarities and differences. Further, a short comparison of the two system architectures will be examined including a side by side assessment of some selected system's hardware mass.

A concept for a miniature, mechanically pumped two-phase cooling loop with high thermal performance was developed. In this feasibility study, a miniature, annular gear pump was inserted into the liquid line of a two-phase LHP-type loop architecture. In contrast to capillary-pumped systems, the functions of liquid pumping and evaporative heat transfer were separated and could be optimized independently. The cooling system was tested in terms of heat transport capability, performance and stability using water as the working fluid. The results show a high heat transfer coefficient of >11 W/(cm2K), a high heat transport capability of >70 W/cm2, and stable working behavior in all orientations. These results were obtained with a device using a simple loop architecture and an evaporator design that was not optimized for this kind of operation.

Future vehicle power systems must achieve greater flexibility and reliability than those used in previous generations. New functions that enhance safety, such as arc detection and wiring integrity verification, are essential for new systems. Embedded autonomous control, and fault correction can be built into Fault Tolerant Processors that integrate into a vehicle Open System Architecture. This approach will provide status and fault detection information to maintenance interfaces and provide fault correction. Safety is enhanced by the prevention of dangerous restarts from crew and personnel. The embedded features allow for pre-flight mission configuration to setup systems before takeoff and on-board and off-board maintenance control. This enables operators to evaluate power system health and history to help reduce turn around time.

The Shuttle orbiter currently uses hydrazine-powered APU’s for powering its hydraulic system pumps. To enhance vehicle safety and reliability, NASA is pursuing an APU upgrade where the hydrazine-powered turbine is replaced by an electric motor pump and battery power supply. This EAPU (Electric APU) upgrade presents several thermal control challenges, most notably the new requirement for moderate temperature control of high-power electronics at 132 °F (55.6 °C). This paper describes how the existing Water Spray Boiler (WSB), which currently cools the hydraulic fluid and APU lubrication oil, is being modified to provide EAPU thermal management.

The Ames Research Center initiated a program in 1975 to provide the critical information required for the design of integrated avionics suitable for general aviation. The program has emphasized the use of data busing, distributed microprocessors, shared electronic displays and data entry devices, innovative low-cost sensors, and improved functional capability. Design considerations include cost, reliability, maintainability, and modularity. As a final step, a demonstration advanced avionics system is being designed, fabricated, and flight tested. The purpose of this paper is to provide a functional description of the Demonstration Advanced Avionics System including a description of the system architecture in order to document the direction that the program is taking.

The design results for a 250 kW switched reluctance aircraft engine starter/generator system power inverter are presented. The starter/generator employs a single switched reluctance machine and a generating system architecture that produces two separate 270 Vdc buses from that single switched reluctance machine. The machine has six phases with three of the phases connected to one inverter supplying 125 kW to one 270 Vdc bus while the other three phases are connected to a second inverter supplying 125kW to the other 270 Vdc bus. Each bus has its own EM1 filter and control in addition to its own inverter. Two types of inverters have been developed, one type employs MOS Controlled Thyristors (MCTs) for the controlled switches and the other type employs Insulated Gate Bipolar Transistors (IGBTs). High-current 500 A peak turn-off MCT modules were specifically developed for the MCT inverters. Two of these modules are placed in parallel to form the required 1000 A switches.

The performance of a more-electric aircraft (MEA) power system electrical accumulator unit (EAU) architecture consisting of a 57000 rpm induction machine (IM) coupled to a controllable shaft load and controlled using direct torque control (DTC) is examined through transient modeling and simulation. The simplicity and extremely fast dynamic torque response of DTC make it an attractive choice for this application. Additionally, the key components required for this EAU system may already exist on certain MEA, therefore allowing the benefits of EAU technology in the power system without incurring a significant weight penalty. Simulation results indicate that this architecture is capable of quickly tracking system bus power steps from full regenerative events to peak load events while maintaining the IM's speed within 5% of its nominal value.

Modeling and Simulation - M&S is recently gaining more importance and emphasis as an essential method for developing engineering systems especially for aerospace and automotive systems, due to their complexity, integration and even human involvement. The main reasons for M&S having that important role nowadays are: 1) M&S can predict system behavior and possible problems. Therefore, it can reduce time and cost for developing systems, it can avoid future corrections into systems, as well. 2) M&S can be used for conception, training, maintenance, etc., requiring less expensive tools and previously preparing people to the real scenario. 3) When it comes to situations that involve aerospace or other products, where high costs are involved, mistakes can be avoided or at least minimized. Summarizing, M&S can reduce project cost and schedule, and improve quality.

The use of control architectures with the Integrated Modular Avionics (IMA) concept (“IMA architectures”) in aerospace and the Integrated Modular Electronics (IME) concept (“IME architectures”) in automotive applications is growing due to its reduced number of hardware such as processors, Line Replaceable Units (LRUs) and Electronic Control Units (ECUs), thereby reducing weight and costs. Furthermore, IMA architectures can perform complex reconfigurations in the case of failures and adapt themselves to changes in network functioning or operating modes, which make a control system very robust. The objective of this work is to discuss the use of an IMA architecture to simulate an aerospace control system responsible for maintaining a vehicle in a predetermined trajectory. To do that, we review the current literature related to IMA architectures and give an overview of their characteristics. Then, we choose an aerospace control system and discuss its simulation using an IMA platform.