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Electrical power system architectures and design concerns for future Mach 3 to 25 aerospace vehicles are discussed. Three conceptual electrical power system architectures were developed to meet the requirements of a future aerospace vehicle. Architectures developed include a hybrid 270 Vdc/115 Vac four- bus architecture, a 270 Vdc four-bus architecture, and a 115 Vac four-bus architecture. The vehicle requirements chosen for this study include electrically driven flight control surface actuators and main engine propellant boost pumps, A trade study was performed to determine the optimum distribution architecture based on total system weight, compatibility of the system with the vehicle loads, component availability, corona susceptibility, development risk, and personnel safety. The trade study identified the hybrid 270 Vdc/115 Vac four bus architecture as the optimum configuration for the given requirements.

The application of Variable speed Constant Frequency (VSCF) electric power generation systems has been limited to aircraft not requiring parallel operation of multiple channels. New applications in large multi-engine aircraft will require continuous parallel operation of installed channels to provide better system reliability and even loading of the generators. Parallel operation of VSCF equipment is similar in implementation to the familiar constant speed generator systems. The VSCF approach yields better overall performance than conventional systems due to inherent accuracies attainable with electronic power systems. Synchronous operation with precise phase angle control between channels is provided during either parallel or split modes of operation. Auto-paralleling of channels thus results in very small amplitude transients and virtually no frequency or phase angle transients on the bus.

The proposed paper will focus on the Air Force's effort that will increase reliability and maintainability and substantially decrease, if not eliminate, the cost of maintaining aircraft battery systems. This effort is recognized throughout the military services, (Army, Navy, and Air Force), and with most of the major aircraft battery manufacturers as the Tiger Team. The Tiger Team began approximately two years ago. The goal of this team is to eliminate battery flight line maintenance shops. The battery that is currently used in most aircraft, vented nickel-cadmium, requires frequent flight line maintenance. This procedure is very costly and time consuming. The proposed paper will overview battery technologies that are being considered for a replacement of the conventional vented nickel-cadmium aircraft battery technology. This overview will also look briefly at the history of aircraft battery systems and how aircraft battery systems arrived at their current status.

Brushless Variable Speed Constant Frequency (VSCF) electric power generation may be obtained using cascaded symmetrically wound machines. The feasibility of using these machines as the basis for a stand-alone aircraft generator system was investigated by the USAF Aero Propulsion Laboratory. The concept is attractive as the system operates without hydraulics and employs a solid state power converter which operates at a fraction of the system output power and frequency. These factors combine to offer a system of relatively low complexity with the potential for high-reliability operation. This paper will discuss the operation of the cascaded doubly fed VSCF generator system and microprocessor control unit.

A highly reliable alternative source for aircraft +28Vdc power is presented. This alternative uses a permanent magnet generator (PMG) and an electronic converter/regulator (C/R). The power system includes such features as independence from the main power system, high power quality with terminal or remote point of regulation (POR), light weight, and constant power availability from engine idle to maximum rpm. Included is a brief system description, a review of steady state and transient performance, and a conclusion with a perspective on future expectations.

The NASA and the military services (NADC and AFWAL) have all promoted the precepts of an “all electric airplane” that would eliminate distributed high pressure hydraulics and multiply different power sources in one airplane. This paper revisits the use of all electric power in military/commercial aircraft and discusses approaches by which the desired implementation might be achieved. The subject and topic of “all electric aircraft” is not new and, in fact, a few U.S. and European aircraft have come very near to justifying that appellation. This paper discusses these past efforts to implement more use of electrics and it reviews the current and emergent technologies that will make the realization of the objective more achievable in the 1990s.

Because of its low density and improved elastic modulus properties, use of aluminum-lithium alloys has good potential for substantial weight savings in aircraft, spacecraft, and missile applications. To study further the possible uses of this material, Douglas Aircraft Company conducted a program to investigate the effects of welding methods on the load-carrying capacities of welded joints made of aluminum-lithium alloy. In the study, laser-welded 2091-T8X aluminum-lithium sheet joints were fabricated and tested. The high-energy intensity and low-heat input generated by laser welding sharply narrowed the heat-affected zone and markedly improved the strength of the weld. Fracture toughness, Jlc, was determined by the J-integral procedure of ASTM E813. The fatigue crack growth rate, da/dN, was measured with precracked specimens and correlated by ΔK. The fracture toughness and fatigue crack growth in the welding zone and in the base metal section were compared.

A variety of components have been successfully forged in mechanically alloyed IncoMAP* Alloy A1-905XL and I/M route Alcan 8090 aluminium-lithium alloys. Open die, ring rolling and closed die forging have been shown to be practical processing routes for both materials. The requirement for cold working 8090 prior to ageing in order that optimum properties are developed has been demonstrated. Components in the IncoMAP material require no heat treatment following forging to develop the required property levels. IncoMAP A1-905XL has been shown to be thermally stable following exposure to temperatures up to 400°C.

Tests with specially instrumented NASA B-737 and B-727 aircraft together with several different ground friction measuring devices have been conducted for a variety of runway surface types and wetness conditions. This effort is part of the Joint FAA/NASA Aircraft/Ground Vehicle Runway Friction Program aimed at obtaining a better understanding of aircraft ground handling performance under adverse weather conditions and defining relationships between aircraft and ground vehicle tire friction measurements. Aircraft braking performance on dry, wet, snow-, and ice-covered runway conditions is discussed together with ground vehicle friction data obtained under similar runway conditions. For a given contaminated runway surface condition, the relationship between ground vehicles and aircraft friction data is identified. The influence of major test parameters on friction measurements such as speed, test tire characteristics, and surface contaminant type are discussed.

Landings of the Space Shuttle Orbiter at 200 knot speeds on the rough, grooved Kennedy Space Center runway have encountered greater than anticipated tire wear, which resulted in limiting landings on that runway to crosswinds of 10 knots or less. The excessive wear stems from wear caused during the initial tire touchdown spin-up. Tire spin-up wear tests have been conducted on a simulated KSC runway surface modified by several different techniques in an effort to reduce spin-up wear while retaining adequate wet cornering coefficients for directional control. The runway surface produced by a concrete smoothing machine using cutters spaced 1 3/4 blades per centimeter was found to give adequate wet cornering while limiting spin-up wear to that experienced in spinups on smooth concrete.

Supersonic Laminar Flow Control (LFC) airplanes with externally braced highly swept LFC wings of high structural aspect ratio (with the sweep increasing towards the wing root) offer particularly high supersonic cruise (L/D)'s with low sonic boom overpressures. At the design cruise condition the flow in the direction normal to the upper surface isobars is transonic (with embedded supersonic zones) and practically shockfree over most of the span. 3-body type supersonic LFC airplanes with a central fuselage and two smaller outboard bodies (alleviating wing bending and torsion) enable further increased spans and aspect ratios to reduce accordingly the lift induced wave- plus vortex drag as well as the volume induced wave drag (L/D)Cruise thus increases further.

The concept of utilizing an electrically actuated aircraft braking system could result in greater fire safety, the elimination of centralized hydraulics, and compatibility with an all-electric aircraft. Using the Air Force A-10 as a test bed, the first fully functional electric brake was laboratory tested, qualified, and installed on an aircraft for testing. On-aircraft testing was curtailed due to a dynamic instability between the brake and landing gear. An extensive laboratory dynamometer test program was substituted. The prototype electric brake demonstrated performance nearly equivalent to the production hydraulic brake with a potential for more accurate torque control.

Most previous studies of the drag of two-dimensional airfoils consider only the total drag. The present report gives results of a study of three airfoils, using the Eppler program, to determine the distribution of friction drag along the chord and to obtain relative values of friction drag and pressure drag over a wide range of angle of attack and Reynolds number. The effects of boundary-layer suction in the turbulent region of the boundary layer of two of the airfoils are also investigated. The pressure drag is found to be an important component of the total drag, reaching values of 60 to 80 percent of the total drag near the stall. The use of suction producing a uniform inflow in the turbulent region of the boundary layer results in large increases in maximum lift, and increases the skin-friction drag but reduces or even changes the sign of the pressure drag.

The present Study sets forth criteria for the design of airport safety areas and land uses therein. The Study has been in preparation and in real-world testing and verification since 1972. Its prototype was developed by the writer for the Airport Land Use Commission (ALUC) of the County of Santa Clara, California, which adopted the criteria established by the Study during 1973. The accident-potential concepts underlying the 1972 Study were also adopted and incorporated by the U.S. Department of Defense in its AICUZ policies and guidelines.

Laminar flow flight experiments conducted over the past fifty years will be reviewed. The emphasis will be on flight testing conducted under the NASA Laminar Flow Control Program which has been directed towards the most challenging technology application- the high subsonic speed transport. The F111/TACT NLF Glove Flight Test, the F-14 Variable Sweep Transition Flight Experiment, the 757 Wing Noise Survey and NLF Glove Flight Test, the NASA Jetstar Leading Edge Flight Test Program, and the recently initiated Hybrid Laminar Flow Control Flight Experiment will be discussed. To place these recent experiences in perspective, earlier important flight tests will first be reviewed to recall the lessons learned at that time.

The NASA Lewis Research Center is presently conducting an aircraft icing research program, the major thrust of which, is to advance technologies that improve our ability to model the icing phenomenon and its effect on aircraft. The approach employs three interrelated elements: analysis; wind tunnel experiments; and, considerable flight testing in natural icing clouds. This paper presents a brief overview of this program with emphasis on recent accomplishments.

Scale model studies of the Shuttle Orbiter Arrestment System have been completed. The system was tested with a 1/27.5 scale model at the NASA Langley Research Center and a 1/8 scale model at All American Engineering Company. The purpose of these studies was to determine the proper net arrestment system configuration to bring the Orbiter to a safe stop in the event of a runway overrun with minimal damage. Tests were conducted for centerline engagements and off-center engagements at simulated speeds up to 95 knots full scale. The results of these studies defined the net-orbiter interaction, corrections to prevent underwing engagements, corrections necessary to prevent net entanglement in the main gear, the dynamics of off-centerline engagements, and the maximum number of vertical straps that might become entangled with the nose gear.

One of the factors needed to describe the wear behavior of the Space Shuttle Orbiter main gear tires is their behavior during the spin-up process. An experimental investigation of tire spin-up processes was conducted at the NASA Langley Research Center's Aircraft Landing Dynamics Facility (ALDF). During the investigation, the influence of various parameters such as forward speed and sink speed on tire spin-up forces were evaluated. A mathematical model was developed to estimate drag forces and spin-up times and is presented. The effect of prerotation was explored and is discussed. Also included is a means of determining the sink speed of the orbiter at touchdown based upon the appearance of the rubber deposits left on the runway during spinup.

During the last two decades, high performance hydraulic systems have been developed, involving ever increasing pressures. The standard pressure in missile and aircraft systems was 3000 p.s.i. during the 1950's. Then pressures were increased to 4500 p.s.i., and now to 8000 p.s.i. These higher pressure systems are being adopted in an attempt to handle higher power and to minimize component weights. Because of this increase in pressure and power, severe dynamic problems have resulted, such as extreme pressure transients, or “water hammer”. This paper describes a time domain digital simulation computer program called RoHDA, which can be used to simulate the dynamic conditions of a system of any complexity. RoHDA is the acronym for Rockwell Hydraulic dynamic Analysis.

The HYdraulic TRansient ANalysis (HYTRAN) program has under gone extensive revisions that correct and enhance the program to meet the needs of those analyzing and/or designing aerospace hydraulic systems. Presented in this paper are a discussion of the revisions, a detailed explanation for the changes, and evidence from flight-test data that the improved HYTRAN program can meet the needs of the aerospace industry. The revisions affect all aspects of the HYTRAN program: steady-state analysis, transient analyses, hydraulic component models, and overall program improvements. The effects of these revisions on hydraulic analyses have been to 1) correct previous steady-state and transient solution errors, 2) reduce user handling time, 3) increase computational efficiency, and 4) improve HYTRAN's hydraulic component library.