Search Results

The main objective of the current paper was to investigate the fluid flow and pressure loss characteristics of DPF substrates with asymmetric channels utilizing 3-D Computational Fluid Dynamics (CFD) methods. The ratio of inlet to outlet channel width is 1.2. First, CFD results of velocity and static pressure distributions inside the inlet and outlet channels are discussed for the baseline case with both forward and reversed exhaust flow. Results were also compared with the regular DPF of same cell structure and wall material properties. It was found that asymmetrical channel design has higher pressure loss. The lowest pressure loss was found for the asymmetrical channel design with smaller inlet channels. Then, the effects of DPF length and filter wall permeability on pressure loss, flow and pressure distributions were investigated.

Shortly after World War II, as aircraft became more sophisticated and power-assist, flight-control functions became a requirement, hydraulic system operating pressures rose from the 1000 psi level to the 3000 psi level found on most aircraft today. Since then, 4000 psi systems have been developed for the U.S. Air Force XB-70 and B-1 bombers and a number of European aircraft including the tornado multirole combat aircraft and the Concorde supersonic transport. The V-22 Osprey incorporates a 5000 psi hydraulic system. The power levels of military aircraft hydraulic systems have continued to rise. This is primarily due to higher aerodynamic loading, combined with the increased hydraulic functions and operations of each new aircraft. At the same time, aircraft structures and wings have been getting smaller and thinner as mission requirements expand. Thus, internal physical space available for plumbing and components continues to decrease.

Shortly after World War II, as aircraft became more sophisticated and power-assist, flight-control functions became a requirement, hydraulic system operating pressures rose from the 1000 psi level to the 3000 psi level found on most aircraft today. Since then, 4000 psi systems have been developed for the U.S. Air Force XB-70 and B-1 bombers and a number of European aircraft including the tornado multirole combat aircraft and the Concorde supersonic transport. The V-22 Osprey incorporates a 5000 psi hydraulic system. The power levels of military aircraft hydraulic systems have continued to rise. This is primarily due to higher aerodynamic loading, combined with the increased hydraulic functions and operations of each new aircraft. At the same time, aircraft structures and wings have been getting smaller and thinner as mission requirements expand. Thus, internal physical space available for plumbing and components continues to decrease.

In 1995, the Naval Research Laboratory (NRL) began a program to investigate whether a 90 Ah dependent pressure vessel (DPV) NiH2 battery pack could be a lower volume replacement for a 90 Ah NiH2 IPV spacecraft battery. Nickel Hydrogen (NiH2) dependent pressure vessel (DPV) battery cells are presumed to offer all the features of the NiH2 IPV battery cell with considerably less volume. To achieve this reduction in volume, the DPV cell utilizes a canteen shaped pressure vessel with reduced thickness wall, flat sides and curved ends. The cells can be packaged similar to prismatic nickel cadmium battery cells. Moreover, like NiCd cells, a fully charged DPV cell must rely upon an adjacent battery cell or structure for support and to maintain pressure vessel integrity. Seventeen 90 Ah NiH2 DPV cells were delivered to NR in 1998 for qualification tests. An eleven-cell half battery pack was manufactured and tested to validate the advantages of the DPV design.

The 912 engine is a well known 4-cylinder horizontally opposed 4-stroke liquid-/air-cooled aircraft engine. The 912 family has a strong track record: 40 000 engines sold / 25 000 still in operation / 5 million flight hours annually. 88% of all light aircraft OEMs use Rotax engines. The 912iS is an evolution of the Rotax 912ULS carbureted engine. The “i” stands for electronic fuel injection which has been developed according to flight standards, providing a better fuel efficiency over the current 912ULS of more than 20% and in a range of 38% to 70% compared to other competitive engines in the light sport, ultra-light aircraft and the general aviation industry. BRP engineers have incorporated several technology enhancements. The fully redundant digital Engine Control Unit (ECU) offers a computer based electronic diagnostic system which makes it easier to diagnose and service the engine.

A low pressure axial fan for benchmarking numerical methods in the field of aerodynamics and aeroacoustics is presented. The generic fan for this benchmark is a typical fan to be used in commercial applications. The design procedure was according to the blade element theory for low solidity fans. A wide range of experimental data is available, including aerodynamic performance of the fan (fan characteristic curve), fluid mechanical quantities on the pressure and suction side from laser Doppler anemometer (LDA) measurements, wall pressure fluctuations in the gap region and sound characteristics on the suction side from sound power and microphone array measurements. The experimental setups are described in detail, as to ease reproducibility of measurement positions. This offers the opportunity of validating aerodynamic and aeroacoustic quantities, obtained from different numerical tools and procedures.

A low pressure axial fan for benchmarking numerical methods in the field of aerodynamics and aeroacoustics is presented. The generic fan for this benchmark is a typical fan to be used in commercial applications. The design procedure was according to the blade element theory for low solidity fans. A wide range of experimental data is available, including aerodynamic performance of the fan (fan characteristic curve), fluid mechanical quantities on the pressure and suction side from laser Doppler anemometer (LDA) measurements, wall pressure fluctuations in the gap region and sound characteristics on the suction side from sound power and microphone array measurements. The experimental setups are described in detail, as to ease reproducibility of measurement positions. This offers the opportunity of validating aerodynamic and aeroacoustic quantities, obtained from different numerical tools and procedures.

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.

This paper describes some of the steps necessary to certificate a camera installation in a Part 23 aircraft. The camera is large and necessitates the severing of a major structural member (keel beam). Damage tolerance analysis is required because of the penetration of the pressure vessel. Cable rerouting is necessary because of the location of the hole. The design, the basis for the structural substantiation and the structural substantiation of this design are discussed in this paper.

Catalytic combustion coupled with activated carbon and molecular sieve adsorbents is applicable to all areas of air and gas clean up ranging from high to low levels of pollutants and trace contaminants control in a spacecraft environment is of no exception. In this study we propose a combined activated charcoal and catalytic combustion system based on a 70 watt power input achieving 350°C, operating on a 6 hour per 24 hour day catalytic cycle with an actual flow of 10.6 l min-1 in a residual free volume of 60 m3.

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 objective of this investigation was to measure and document the existence of any significant differences in physical performance under operational conditions between the Launch Entry Suit (LES) and the new Advanced Crew Escape Suit (ACES). The LES is a partial pressure suit currently worn by astronauts during the launch and entry phases of Shuttle missions. The ACES is a full pressure suit under consideration as a replacement for the LES. One prototype ACES has been fabricated and was used in this investigation. This report presents the results of three tests conducted with six subjects to allow a comparative evaluation of the two suits. The three tests included a reach envelope test, a strength test, and a treadmill test. The reach envelope test measured and compared the maximum hand displacements during horizontal and vertical reaches of both left and right arms in one-g conditions.

This paper describes a computational technique for prediction of the flow and thermal environment in the Space Shuttle Solid Rocket Motor field joint cavities. The SRM field joint hardware has been tested with a defect in the insulation. Due to this defect, the O-ring gland cavities are pressurized during the early part of the ignition. A computer model has been developed to predict the flow and thermal environment through the simulated flaw, during the pressurization of the field joint. The transient mass, momentum, and energy conservation equations in the flow passage in conjunction with the thermodynamic equation of state are solved by a fully implicit iterative numerical procedure. Since this is a conjugate flow and heat transfer problem, wall temperatures are calculated by solving the one-dimensional transient heat conduction equation in the solid along with the other governing equations. The pressure and temperature predictions have been compared with the test data.

This paper develops a set of seven general equations which describe various portions of a standard engine performance chart. A regression analysis technique is then employed to evaluate the constants in these equations which identifies them to a particular engine. Computer programs which employ this technique are provided. A final program is provided which utilizes the equations to yield a value of brake horsepower for an input of any combination of RPM, manifold air pressure, air temperature, and pressure altitude.

This paper reports the results of a program to develop a pneumatic actuation system for an experimental ejection seat trainer. The device was developed under contract to the U. S. Naval training Device Center, Port Washington, New York. A method of pneumatic open loop flow and pressure control is presented which results in predictable load dynamics during the ejection process. Close control of maximum rate of acceleration and maximum acceleration is obtained for a wide range of load weights. System analysis and individual component designs are discussed. Test data and calibration procedures are also presented. Although the present device is highly specialized, the design methods described in this paper are applicable to any high speed pneumatic ejection system requiring close control of load dynamics.

The considerable progress made by turbo-machinery design in the last decade has been paced by the rigorous demands of the customers and the competitive pressures of the market place. The requirements have been for significant improvements in product operability, performance, cost, reliability, durability, maintainability and weight. Four inter-related fronts have been responsible for much of this progress: ADVANCEMENTS IN MECHANICAL DESIGN CONFIGURATIONS - such as integral blades and disks, more efficient hot-part cooling, sophisticated clearance control systems and welded rotors -have afforded improvements in virtually all measurements of merit. ADVANCED MATERIALS AND PROCESSING yielding improved temperature, strength and life properties - have permitted designs with higher cycle pressures, temperatures and tip speeds.

The use of external jets issuing normal to the surface of a body for aerodynamic control has received attention in the past due to experimental observations of favorable interactions between the free stream flow and the jet wake which, in effect, augment the force produced by the jet. The purpose of this study was to perform preliminary trades to determine the effect of variables such as reaction jet location, free stream Mach number, reaction jet stagnation pressure, and reaction jet exit Mach number on augmentation factors for a representative lifting body cruise missile configuration. A computational fluid dynamic analysis, solving the Euler equations, was used to perform the trades listed above. It is shown that the reaction jet augmentation factors are greater, and hence the reaction jet is more effective, at higher free stream dynamic pressures in the subsonic regime.

Gas-permeable membranes that continuously transfer carbon dioxide (CO2) from air to water were investigated in an effort to bypass the operational limitations of expendable solid absorbents currently used for CO2 control in closed-circuit underwater breathing apparatus (UBA). Rebreather UBA CO2 control requirements and known membrane properties were used to create a functional hierarchy of membrane types and CO2 transfer mechanisms, from which one membrane configuration was selected for evaluation. This Direct Contact Membrane Separator (DCMS) employs microporous hydrophobic Hollow Fiber Membrane (HFM) modules to create large membrane areas in small volumes for air-water phase contact without intermixing. Since the micropores in the hydrophobic walls of the hollow fibers are air-filled, gas permeation rates through this membrane are far higher than for any solid or liquid membrane.

A dynamic simulation model has been developed for a vapor-cycle cooling system designed for aircraft applications using the latest technology developments. The heat exchanger models use multiple-, lumped-parameter, fixed-length elements based on coupled thermal and mass storage effects, and flow equations that incorporate the effects of thermal expansion and contraction. This model is developed to include the two-phase constant pressure temperature gradient unique to refrigerant mixtures. The full system model incorporates global mass conservation which is essential for accurate pressure levels and, thus, dynamic response and steady state performance. Phase boundary-based coordinate transformations on the nonazeotropic refrigerant mixture property data result in improved accuracy and computation efficiency. The simulation is developed with modular components with causality defined to minimize connection states and thus execution time.

A discussion is given of the ongoing research related to laminar flow airfoils, nacelles, and wings where the laminar flow is maintained by a favorable pressure gradient, surface suction or a combination of the two. Design methologies for natural laminar flow airfoil sections and wings for both low and high speed applications are outlined. Tests of a 7-foot chord, 23° sweep laminar-flow-control-airfoil at high subsonic Mach numbers are described along with the associated stability theory used to design the suction system. The state-of-the-art of stability theory is simply stated and a typical calculation illustrated. In addition recent computer simulations of transition using the time dependent Navier-Stokes (N-S) equations are briefly described. Advances in wind tunnel capabilities and instrumentation will be reviewed followed by the presentation of a few results from both wind tunnels and flight. Finally, some suggestions for future work will complete the paper.