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This Life Support Laboratory consists of a simulator of the spacecraft called Nautilus, which houses Air Revitalization Subsystem, Atmospheric Control and Supply, and Fire Detection and Suppression in the Equipment Area. There are supporting facilities including a Human Metabolic Simulator, simulated Low and Moderate Temperature Coolant Loop, chemical analysis bench, purified water supply, vacuum and gas supplies. These facilities are scheduled to be completed and start to operate for demonstration purposes by March 2005. There are an ARES Ground Model (AGM) and a Trace Contaminant Control Assembly in the ARS. The latter will be integrated with the AGM and a Condensing Heat Exchanger. The unit of AGM is being engineered, built, and will be delivered in early 2005 by EADS Space Division. These assemblies will be operated for sensitivity analysis, integration and optimization studies. The main goal is the achievement for optimal recovery of oxygen.

eROSITA (extended ROentgen Survey with an Imaging Telescope Array) is a powerful X-ray telescope under development by the Max-Planck-Institut für extraterrestrische Physik (MPE) in Garching, Germany. eROSITA is the core instrument on the Russian SRG1 mission which is planned for launch in 2011. It comprises seven nested Wolter-I grazing incidence telescopes, each equipped with its own CCD camera. The mirror modules have to be maintained at 20°C while the cameras are operated at -80°C. Both, mirrors and CCDs have to be kept within tight limits. The CCD cooling system consists of passive thermal control components only: two radiators, variable conductance heat pipes (VCHP) and two special thermal storage units. The orbit scenario imposes severe challenges on the thermal control system and also on the attitude control system.

Future space exploration missions will require a lightweight spacesuit that expends no consumables. This paper describes the design and performance of a prototype heat rejection system that weighs less than current systems and vents zero water. The system uses regenerable LiCl/water absorption cooling. Absorption cooling boosts the heat absorbed from the crew member to a high temperature for rejection to space from a compact, non-venting radiator. The system is regenerated by heating to 100°C for two hours. The system provides refrigeration at 17°C and rejects heat at temperatures greater than 50°C. The overall cooling capacity is over 100 W-hr/kg.

The X-Wing concept employs a single lifting system for all modes of flight. The lifting system is comprised of four very rigid, circulation control wings with blowing for lift modulation and control. For hover and low speed flight, the wings rotate such as the rotor of a helicopter. For high speed flight, the wings are stopped in an “X” configuration across the fuselage - from which the name of the concept is derived - with two forward-swept wings and two aft-swept wings. Such a vehicle is also envisioned to have an integrated gas turbine propulsive system for all flight modes. At low speeds, the gas generators) would drive a shaft to turn the wings and the circulation control compressor as well as a set of propulsive fans. For high-speed flight, the shaft would drive only the compressor and accessories as the fans propel the vehicle. The X-Wing concept has been underdevelopment for over 15 years.

The substantiation of heat pipe usage in passive radiative cooling systems on temperature level (190…240) K for space optical sensors is presented. Heat pipes can be sound practice like heat conducting lines between sensor and radiator particularly at distances more 0.2 m and irreplaceable at distances (0.5…2) m. Embedding heat pipe with radiator allows to create the uniform temperature basis in case of several sensors connection to single radiator and to improve radiator efficiency. It is analyzed approach to design of thermocontrol and cooling radiative systems with heat pipes to reduce sensitiveness to external light disturbances and to enlarge area of radiative system application. The results of design, thermovacuum test and flight operation of thermocontrol radiative system samples are under discussion as well.

Selection of working fluid is one of the main criterions for designing of heat pipes thermal control systems (TCS) for space application. In this paper we will describe how we solved the task of development of the TCS with working fluid of high thermal physical properties. In 2004-2006 we developed the Engineering model of Deployable Radiator based on Loop Heat Pipe by CAST purchase order. It was developed for qualification tests. Ammonia application as LHP working fluid is stipulated by its high thermal physical properties. However Ammonia freezing temperature is of minus 77ºC. Such fact impedes Ammonia application when operation temperatures of LHP Radiator are lower than this value, for example, It takes several tens of hours to orbit a spacecraft and prepare it for work (at that moment the spacecraft is out of power supply) and the working fluid can be frozen in a condenser-radiator when the spacecraft being in the shadow over a long period of time.

Spacecraft radiators reject heat to their surroundings. Radiators can be deployable or mounted on the body of the spacecraft. NASA's Crew Exploration Vehicle is to use body mounted radiators. Coatings play an important role in heat rejection. The coatings provide the radiator surface with the desired optical properties of low solar absorptance and high infrared emittance. These specialized surfaces are applied to the radiator panel in a number of ways, including conventional spraying, plasma spraying, or as an appliqué. Not specifically designed for a weathering environment, little is known about the durability of conventional paints, coatings, and appliqués upon exposure to weathering and subsequent exposure to solar wind and ultraviolet radiation exposure. In addition to maintaining their desired optical properties, the coatings must also continue to adhere to the underlying radiator panel.

An analysis is presented of the relative suitability of sodium, potassium, rubidium, and cesium as working fluids in a high temperature, Rankine Cycle, space power plant. Turbine inlet temperatures of from 1800 to 2000 F with corresponding condensing temperatures of from 1240 to 1530 F are considered. The criteria by which the fluids are evaluated are the thermodynamic cycle characteristics, heat transfer and fluid friction characteristics, metallurgical compatibility, and the influence of the fluids on the design of the turbine, bearings, radiator, generator, and pump. The turbogenerator unit is thought to be the most critical component and it is found that the working fluid will determine the required number of turbine stages and will therefore establish the turbogenerator bearing arrangement. It is not known whether blade erosion will be a problem.

Vought has participated with the Navy and NASA in studies of various subsonic and supersonic V/STOL aircraft. From these studies, a unique propulsion concept, the Tandem Fan, has been developed. A. 11 scale subsonic V/STOL Tandem Fan model has recently been tested in Vought's large ground effects facility (LGEFF) and Vought's 7 × 10 ft low speed wind tunnel. The Vought large ground effects facility, provides V/STOL testing at hover in the presence of 0 to 30 knots winds in a free circulation environment. The Tandem Fan tests in this facility consisted of over 100 runs to evaluate both attitude and control effects. Individual fan thrust modulation was tested to evaluate pitch and roll control effectiveness, and exhaust nozzle perturbations were tested for yaw control. The model was then installed in the Vought 7 × 10 ft. low speed wind tunnel to evaluate the cruise and upper transition flight regions.

This paper describes a class of “powered lift” aircraft called volantors that utilize multiple ducted fans for generating thrust. Thrust is directed vertically for lift during vertical takeoff and landing and horizontally during cruise. The change in thrust direction is accomplished by partially rotating the ducted fans and partially redirecting the airflow at the duct exit through a series of variable camber vanes. This arrangement avoids the potential of duct leading edge stall during transition that arises if the thrust is re-directed solely by rotating the ducts through 90°. In addition to the above fundamental design arrangement this paper discusses the technical issues and developments that need to be addressed in order to provide a volantor that is practical enough for large scale personal use.

A new and advanced portable life support system (PLSS) for space suit surface exploration will require a durable, compact, and energy efficient system to transport the ventilation stream through the space suit. Current space suits used by NASA circulate the ventilation stream via a ball-bearing supported centrifugal fan. As NASA enters the design phase for the next generation PLSS, it is necessary to evaluate available technologies to determine what improvements can be made in mass, volume, power, and reliability for a ventilation transport system. Several air movement devices already designed for commercial, military, and space applications are optimized in these areas and could be adapted for EVA use. This paper summarizes the efforts to identify and compare the latest fan and bearing technologies to determine candidates for the next generation PLSS.

Engine manufacturers recognize that either a variable pitch fan blade or a variable area fan nozzle can provide the necessary fan stall/flutter margin to take the next step in engine cycle improvement with enhanced propulsive efficiency and low noise level. The patented Rotating Composite Technologies, LLC variable pitch fan design uniquely offers the same or lower fan hub/tip radius ratio of conventional fixed pitch designs and, with its light weight composite fan blades integrated with an even smaller disk, offers weight savings as well. A variable fan nozzle conversely adds additional weight to the propulsion system and notably will complicate aircraft installation issues relative to competitive manufacturer's engines.

Variable cycle engines (VCEs) hold promise as an enabling technology for supersonic business jet (SBJ) applications. Fuel consumption can potentially be minimized by modulating the engine cycle between the subsonic and supersonic phases of flight. The additional flexibility may also contribute toward meeting takeoff and landing noise and emissions requirements. Several different concepts have been and are currently being investigated to achieve variable cycle operation. The core-driven fan stage (CDFS) variable cycle engine is perhaps the most mature concept since an engine of this type flew in the USAF Advanced Tactical Fighter prototype program in the 1990s. Therefore, this type of VCE is of particular interest for potential commercial application. To investigate the potential benefits of a CDFS variable cycle engine, a parametric model is developed using the NASA Numerical Propulsion System Simulation (NPSS).

The thermal control of lunar landers/rovers necessitates the use of a system to allow heat rejection to the high temperature lunar environment. In this context a vapour compression heat pump which is a proven technology in terrestrial and aeronautical applications has been studied; its suitability in providing 2 kW cooling capability with adequate temperature lift for final heat rejection by space radiators is assessed. The stringent requirements of space-based hardware in terms of temperature lift, compactness, mass, performance and reliability necessitates optimization studies. Mass optimization of the heat pump components has been carried out, as well as selection of refrigerants and thermodynamic cycles most suited for the application.

Spacecraft cold-plate designs are small, relatively complex, and expensive. A large-size, low-cost approach is required for the new generation of space power modules and platforms. This paper describes part of an on-going investigation of the feasibility of using large aluminum extrusions attached side-by-side to form a large, two-sided cold plate with integral coolant passages. The objective was to establish and verify a cold-plate-to-equipment interface design that would meet thermal performance requirements and allow easy replacement by extravehicular activity (EVA) component changeout. Approximately twelve filler materials and six interface designs were studied, and these were selected: three filler approaches, an existing adjustable interface pressure design, and a conventional design with no filler and no adjustment. Thermal conductance tests performed on simulated implementations of these approaches measured the contact conductances of each.

The United States Navy F/A-18 aircraft uses two 30/40 KVA generating systems to provide precise 400 Hz power to their respective isolated busses. Each generating system is composed of a six phase generator running at a speed proportional to engine rpm, feeding an SCR-based, naturally commutated cycloconverter. This was the first integral package with shared cooling oil, production, 400 Hz VSCF system. Over 2,000 units have been produced to date. Due to the radical shift from historical, mechanically supplied constant speed technology, the F/A-18 VSCF design initially raised numerous reliability questions. This paper serves to address those concerns and provide development questions with historical field performance/analyses in response. Reliability predictions using MIL-HDBK-217 procedures are compared to Navy 3-M field performance data over the last eight year production period.

Noise environmental control is an important design consideration for Space Station Freedom (SSF), both for crew safety and productivity. Acoustic noise requirements are established to eliminate fatigue and potential hearing loss by crew members from long term exposure and to facilitate speech communication. VAPEPS (VibroAcoustic Payload Environment Prediction System) is currently being applied to SSF for prediction of the on-orbit noise and vibration environments induced in the 50 to 10,000 Hz frequency range. Various sources such as fans, pumps, centrifuges, exercise equipment, and other mechanical devices are used in the analysis. The predictions will be used in design tradeoff studies and to provide confidence that requirements will be met. Preliminary predictions show that the required levels will be exceeded unless substantial noise control measures are incorporated in the SSF design.

High energy density regenerative fuel cell systems that are used for energy storage require novel approaches to integrating components in order to preserve mass and volume. A lightweight Unitized Regenerative Fuel Cell (URFC) Energy Storage System concept is being developed at the NASA Glenn Research Center (GRC). This Unitized Regenerative Fuel Cell System (URFCS) minimizes mass by using the surface area of the hydrogen and oxygen storage tanks as radiating heat surfaces for overall thermal control of the system. The waste heat generated by the URFC stack during charging and discharging is transferred from the cell stack to the surface of each tank by loop heat pipes, which are coiled around each tank and covered with a thin layer of thermally conductive carbon composite. The thin layer of carbon composite acts as a fin structure that spreads the heat away from the heat pipe and across the entire tank surface.

Current state-of-the-art space radiators are too heavy (5-7 kg/m2) and voluminous to be feasible for some future space missions. Accordingly, there is a need for a revolutionary advanced space radiator system. This paper describes an effort to satisfy that need through the development of an unfurlable heat pipe radiating system. The innovative portion of the unfurlable radiator is the pressure envelope: it is a thin, flexible, heat sealable polymer/metal laminate that is vacuum tight. The laminate allows the radiator to be compactly rolled or folded, easily stowed for transit to space and then unfurled to present a large radiating surface. Condensate return to the evaporator is achieved by a combination of capillary pumping via a flexible porous cable wick and the advanced capillary pumped loop methods of entrainment. The mass density of the radiator is 1.76 kg/m2, representing a reduction in mass of at least a factor of 3 over current radiator technology.

A proof-of-concept (POC) lightweight lunar radiator was fabricated and tested. The POC radiator has a specific weight of 5 kg/kW one quarter the specific weight of current space radiators. The significant weight reduction was due to the radiator's unique design. It is a multicellular heat pipe radiator utilizing the lunar gravity for condensate return. The innovation of this radiator is the laminated film material used as the heat pipe envelope. By utilizing a flexible, durable, leak tight laminate structure instead of the typical ridged heat pipe envelope, significant weight reductions were achieved. In addition, the resulting radiator is extremely flexible, allowing it to be rolled or folded and compactly stored during transit to the lunar surface. Testing demonstrated that a laminated film heat pipe radiator offers improved performance and significant mass savings over conventional space radiators.