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02 AERODYNAMICS 20000061443 Synaps Ingenieur-Gesellschaft m.b.H., Bremen, Germany A CONCEPTUAL DESIGN METHODOLOGY TO PREDICT THE WAVE DRAG OF A TRANSONIC WING Kribler, T., Synaps Ingenieur-Gesellschaft m.b.H., Germany; Aerodynamic Design and Optimisation of Flight Vehicles in a Concurrent Multi-Disciplinary Environment; June 2000, pp. 27-1 - 27-8; In English; See also 20000061419; Copyright Waived; Avail: CASI; A02, Hardcopy A conceptual design methodology to predict the wave drag of a transonic wing for use within multidisciplinary aircraft design was developed. To achieve this, a database of cross section designs optimized with respect to total drag was set up varying the design parameters Ma, t/c, C(sub L) and Re. Mathematical formulations for the aerodynamic cross section characteristics total drag, viscous drag and the local shock location were derived from the database as functions of the design parameters. The cross section wave drag was then derived using these formulations. A locally infinite swept wing is assumed and simple sweep theory using the shock sweep angle is used to transform the wave drag. The wave drag of a 3-D wing is predicted summing locally infinite swept wing sections in spanwise direction. The achieved drag prediction is accurate enough for use within conceptual aircraft design and predicts well the trends in wave drag development as a function of the design parameters Ma, t/c, C(sub L), Re and the wing planform. Author Aircraft Design; Design Analysis; Wave Drag; Transonic Flow; Wing Planforms 20000061444 Florida Univ., Dept. of Aerospace Engineering, Mechanics and Engineering Science, Gainesville, FL USA AIRFOIL AND WING PLANFORM OPTIMIZATION FOR MICRO AIR VEHICLES Sloan, J. G., Florida Univ., USA; Shyy, W., Florida Univ., USA; Haftka, R. T., Florida Univ., USA; Aerodynamic Design and Optimisation of Flight Vehicles in a Concurrent Multi-Disciplinary Environment; June 2000, pp. 28-1 - 28-14; In English; See also 20000061419; Copyright Waived; Avail: CASI; A03, Hardcopy Low Reynolds number flight for micro air vehicles (microAVs) suffers from laminar separation resulting in reduced lift and increased drag. The objective of the present work is to use the response surface methodology (RSM) to identify correlations between the airfoil and the wing planform to facilitate a two-level optimization procedure in which an optimized airfoil and wing planform are reached simultaneously. Several approaches have been considered in this work. A constant cross-section wing is modeled with maximum camber, y(sub c), maximum thickness, y(sub t) and aspect ratio,AR, as design variables at two different Reynolds numbers of 8.0 x 10(exp 4) and 2.0 x 10(exp 5). This is done to determine how the optimal airfoil may change for different aspect ratios and Reynolds numbers. A variable cross-section wing defined by root camber and angle-of-attack and tip camber and angle-ofattack is modeled in order to determine how the optimal airfoil may change from the root to the tip of the wing. Due to the size restrictions on microAVs, a fixed-span approach is used to model an aircraft subject to the constraints of steady flight with the aspect ratio and camber as design variables. This third approach balances trade-offs between wing area, aspect ratio, and Reynolds number in determining the overall flight efficiency. Optimal airfoils exhibit characteristics which change little with wing aspect ratio or location on the wing planform. There appears to be a trend of increasing optimal camber with decreasing Reynolds number. While the optimal design seems to favor airfoils with minimum thickness and relatively modest camber of about 4 to 5% of the chord, a higher camber may be a better choice if higher lift coefficient at minimum power is used as a design goal. Measurements of both the global and the local response surface prediction accuracy combined with design space refinement help to assess the reliability of the response surface approximations and optimal design predictions. Author Airfoils; Wing Planforms; Optimization; Camber; Design Analysis; Low Reynolds Number 20000061446 Aerospatiale, Matra Missiles, Chatillon, France MULTI-FLIGHT CONDITION OPTIMIZATION OF THREE DIMEN- SIONAL SUPERSONIC INLETS Carrier, Gerald, Aerospatiale, France; Bourdeau, Christophe, 8 Aerospatiale, France; Knight, Doyle, Rutgers Univ., USA; Kergaravat, Yan, Aerospatiale, France; Montazel, Xavier, Aerospatiale, France; Aerodynamic Design and Optimisation of Flight Vehicles in a Concurrent Multi-Disciplinary Environment; June 2000, pp. 30-1 - 30-10; In English; See also 20000061419 Contract(s)/Grant(s): DDM980001N; Copyright Waived; Avail: CASI; A02, Hardcopy This paper presents an innovative methodology to address the three-dimensional supersonic inlet design problem. An efficient and robust process allows to optimize the aerodynamic performance of inlets for multiple flight conditions. This optimization process links together an optimizer with a fast and accurate simulation tool into an automated optimization loop. The implementation of this new design technique and its applications to two different test cases are presented, namely, the optimization for a single cruise condition, and the optimization for a mission comprised of acceleration, cruise and maneuver phases. The mission-optimized inlet achieves better overall performance than the cruise-optimized inlet. Author Flight Conditions; Supersonic Inlets; Procedures; Three Dimensional Models 20000108802 Georgia Inst. of Tech., Aerospace Systems Design Lab., Atlanta, GA USA TECHNOLOGIES FOR FUTURE PRECISION STRIKE MISSILE SYSTEMS: MISSILE AEROMECHANICS TECHNOLOGY Fleeman, Eugene L., Georgia Inst. of Tech., USA; Technologies for Future Precision Strike Missile Systems; September 2000, pp. 2-1 - 2-10; In English; See also 20000108801; Original contains color illustrations; Copyright Waived; Avail: CASI; A02, Hardcopy This paper provides an assessment of the state-of-the-art of new aeromechanics technologies for future precision strike missile systems. The aeromechanics technologies are grouped into specific discussion areas of aerodynamics, propulsion, and airframe materials technologies. Technologies that are addressed in this paper are: Missile aerodynamics technologies- Assessments include aerodynamic configuration shaping, lattice tail control, split canard control, forward swept surfaces, bank-to-turn maneuvering, and flight trajectory shaping; Missile propulsion technologies- Assessments include supersonic air breathing propulsion, high temperature combustors, low drag ramjet inlets, ramjet inlet/airframe integration, high density fuels, and rocket motor thrust magnitude control; and Missile airframe materials technologies- Assessments include hypersonic structure materials, composite structure materials, hypersonic insulation materials, multi-spectral domes, and reduced parts count structure. Author Missile Systems; Aerodynamics; Propulsion; Missile Design; Airframes; Missile Components; Missile Configurations; Missile Control; Radar Homing Missiles 20010009842 National Aerospace Lab., Amsterdam, Netherlands DATA FROM AGARD REPORT 702: NACA 64A006 OSCILLAT- ING FLAP; NACA 012 OSCILLATORY AND TRANSIENT PITCH- ING; NLR 7301 SUPERCRITICAL AIRFOIL OSCILLATORY PITCHING AND OSCILLATING FLAP; AND ZKP WING, OSCIL- LATING AILERON Landon, R. H., Aircraft Research Association Ltd., UK; Verification and Validation Data for Computational Unsteady Aerodynamics; October 2000, pp. 29-96; In English; See also 20010009839; Copyright Waived; Avail: CASI; A04, Hardcopy In the late seventies a need was perceived for standard comparison cases and experimental data to aid the comparison and validation of the theoretical methods then emerging for unsteady aerodynamics. A Working Group of the AGARD Structures and Materials panel chose a set of 2-D and 3-D configurations and for each configuration defined a set of test cases, including a priority subset, to be used for comparisons. These test cases were fully identified. The chosen configurations were known as the AGARD Aeroelastic Configurations and the chosen cases were denoted as Computational Test (CT) cases. Some of the CT cases were entirely theoretical while others were also the subject of unsteady measurements. The next step undertaken to aid the methods development was to produce an experimental data compendium (AGARD Report 702, which was conceived with the idea of bringing together the experimental data most important for the comparisons. The report was followed by an Addendum, which introduced two additional 3-D
experiments. These reports established an admirable common base for providing experimental data and their value has been demonstrated by the repeated use of the test cases for the entire period since publication. The report has served as a model for the present new compendium of experimental data. It was decided that some of the data cases in the original Report 702 should be reproduced in this document in order to provide more complete coverage in this report with the additional bonus of making available the original data in electronic form to facilitate its continued use to validate calculations. Author Data Acquisition; Flapping; Oscillations; Proving 20010009856 Institute for Aerospace Research, Ottawa, Ontario Canada SELECTED DATA SET FROM STATIC AND ROLLING EXPERI- MENTS ON A 65 DEG DELTA WING AT HIGH INCIDENCE Huang, X. Z., Institute for Aerospace Research, Canada; Lui, T. C., Institute for Aerospace Research, Canada; Hanff, E. S., Institute for Aerospace Research, Canada; Verification and Validation Data for Computational Unsteady Aerodynamics; October 2000, pp. 383-405; In English; See also 20010009839; Copyright Waived; Avail: CASI; A03, Hardcopy This data set is selected from an extensive set of experimental results obtained for configurations with a 65 deg delta wing under static as well as large-amplitude high-rate rolling or pitching conditions at high incidence. The experiments were performed under a joint research program on ‘Non-Linear Aerodynamics under Dynamic Maneuvers’ by the National Research Council of Canada (NRC (IAR)), the U.S. Air Force (USAF (AFOSR, AFRL)) and the Canadian Dept. for National Defence (DND). NASA Ames informally participated in the program through its substantial CFD work on specific test conditions. The experimental results provide both detail pressure measurements and a wide range of flow conditions covering from simple attached flow, through fully developed vortex and vortex burst flow, up to fully-stalled flow at very high incidence. Since this data set includes different levels of physical difficulty, the computational researchers working in unsteady aerodynamics can use it as a staircase approach to the problem of validating their corresponding code. Four schematic and representative configurations# were selected in the experiments: 1) 65 deg delta wing; 2) 80/65 deg double delta wing; 3) 65 deg delta wing with a single vertical tail and a circular ogive forebody, 4) 65 deg delta wing with a single vertical tail and an elliptical cross section forebody whose major axis could be installed either horizontally or vertically. Experiments with the above models include the following test parameters: 1) motion variables (rolling or pitching), 2) modes (static or dynamic), 3) motion waveform (harmonic, ramp-and-hold, free-to-roll and ‘forced’ free-to-roll), 4) observed variables (flow visualization, motion history, steady and unsteady loads and surface pressure), 5) wind tunnel interference assessment (by repeat tests in different wind tunnels), 6) support interference assessment (by repeat tests with different supports). The words of ‘forced’ free-to-roll refer to the experiments performed in the forced mode with the same motion as observed under free-to-roll condition so that the unsteady surface pressures prevailing during free-to-roll motions could be obtained. The installation and support arrangements in the two wind tunnels are shown. The models, rolling rig and pitching rig were designed by IAR. Experiments were conducted both at the IAR and AFRL wind tunnels (LSWT and SARL respectively) summarizes the test matrix. A complete list of tests with corresponding conditions can be found. The comparisons of repeat tests conducted in different wind tunnels and supports confirm that both wind-tunnel as well as support interference are negligible. Due to large number of tests conducted, this data set contains only ten typical cases for the 65 deg delta wing, are shown. These cases were selected to cover typical sets of tests such as static tests and harmonic, ramp- and-hold, free-to-roll and ‘forced’ free-to-roll dynamic tests. Seven spanwise-distributed surface pressure transducers on the up surface of the port wing were used to measure the instantaneous surface pressure during the motion. Three typical sting angles: sigma = 15 deg, 30 deg and 35 deg were selected as being representative of different leading- edge vortex behavior. In the absence of sideslip, at sigma=15 deg the leading-edge vortex is intact over the full length of the model, leading to small non-linearities and time dependence; at gigma=30 deg vortex breakdown occurs over the aft part of the wing leading to severe non-linearities and time dependence; and finally, at gigma=35 AERODYNAMICS 02 deg vortex breakdown is present over the forward portion of the wing resulting in different characteristics. Author Data Acquisition; Delta Wings; Dynamic Tests; Computational Fluid Dynamics; Stabilizers (Fluid Dynamics); Unsteady Aerodynamics 20010009858 Duits-Nederlandse Windtunnel, Brunswick, Germany OSCILLATING 65 DEG DELTA WING, EXPERIMENTAL Lieser, Thomas, Duits-Nederlandse Windtunnel, Germany; Verification and Validation Data for Computational Unsteady Aerodynamics; October 2000, pp. 415-430; In English; See also 20010009839; Copyright Waived; Avail: CASI; A03, Hardcopy This data set contains force and pressure data resulting from static and dynamic measurements on a sharp-edged cropped delta wing with a leading edge sweep of 65# oscillating in different modes. Motivation for the experiment were the provision of experimental data for validation of unsteady computational codes and understanding of the flow past an oscillating delta wing. The model geometry is identical to a geometry used in the Vortex Flow Experiment for Computer Code Validation (VFE), a multinational cooperation which provided experimental data of delta wing configurations in the mid eighties. The geometry of the wing is also used for steady and unsteady calculations within the Western European Armament Group (WEAG, formerly IEPG) TA - 15. The experiments have been performed in 1994 (force measurements) and 1995 (pressure measurements). They were performed in the German-Dutch wind tunnel DNW-NWB at low speeds, the model undergoing pitching, yawing or rolling motions about wind-fixed axes. The choice of the mean angles of attack was closely related to the expected flow types: alpha(sub 0) = 0 deg: In this case the vortex formation will alternate between the upper and the lower surface of the configuration during the pitching motion. Alpha(sub 0) = 9 deg: Vortices will be present over the upper surface of the configuration and no vortex breakdown will occur during the whole cycle of the pitching motion. Alpha(sub 0) = 15 deg and alpha(sub 0) = 21 deg: These conditions are related to mixed cases without vortex breakdown over the configuration at low angles of attack and with vortex breakdown at high angles of attack during the cycle of motion. Alpha(sub 0) = 27 deg: Vortices with vortex breakdown are expected to occur over the upper surface of the configuration and this type of flow will be present during the whole cycle of the pitching motion. Alpha(sub 0) = 42 deg: During the cycle of the pitching motion the flow is expected to switch between a vortex-type flow with vortex breakdown and a dead-water-type flow. Alpha(sub 0) = 48 deg: In this case a dead-water-type flow is expected during the whole cycle of motion. The mean angles of attack at which no vortex breakdown occurs during the complete cycle of motion are simpler to treat numerically. Therefore the pitching oscillations about alpha(sub 0) = 9 deg was the first case to be included in a WEAG TA-15 common exercise. The other case included in that common exercise is the pitching oscillation about alpha(sub 0) = 21 deg, the reduced frequency being 0.56 for all mean angles of attack. Results from unsteady Euler and Navier-Stokes calculations of pitching oscillation about alpha(sub 0) = 9 deg with an amplitude of Delta(alpha)(sub 0) = 3 deg by W. Author Aerodynamic Configurations; Computational Fluid Dynamics; Delta Wings; Vortex Breakdown; Pressure Measurement; Oscillating Flow 20010009859 DaimlerChrysler Aerospace A.G., Military Aircraft, Munich, Germany OSCILLATING 65 DEG DELTA WING, NUMERICAL Fritz, Willy, DaimlerChrysler Aerospace A.G., Germany; Verification and Validation Data for Computational Unsteady Aerodynamics; October 2000, pp. 431-435; In English; See also 20010009839; Copyright Waived; Avail: CASI; A02, Hardcopy This data set consists of steady and unsteady numerical solutions of a sharp-edged cropped delta wing with a leading edge sweep of 65 deg undergoing a pitching oscillation. The geometry of the wing corresponds with the geometry of the wind tunnel model described in the previous data set, the difference being the absence of the fuselage in the numerical model. The presence of the fuselage on the upper surface flow is believed to have an effect at small angles of attack only on the forward region of the wing and to have an effect on the location of vortex breakdown at large angles of attack. The pitching oscillation has an amplitude of 3 deg, the mean angle of attack is 9 deg. The position of the oscillation axis and the 9
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Page 5 and 6: 93 SPACE RADIATION N.A. Includes co
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Page 9 and 10: inging the smart technologies to re
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20000020800 Rolls-Royce Ltd., Milit
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which recently designed a test rig
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namics, heat transfer, and thermoel
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aircraft in use are investigated. F
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such as blades and vanes. The need
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20010067729 Georgia Inst. of Tech.,
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20010067749 California Univ., Combu
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19990053163 Academy of Sciences of
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the development of airframe and lan
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esults of steady and unsteady compo
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Aerospatiales, France; Active Contr
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degree of instability of the unaugm
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efficacy. All fourteen pilots ident
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Alpha. This report presents some im
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combustors have been a serious prob
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BMW-RR was investigated with optica
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Ingenieurbuero fuer Thermoakustik G
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autoignition conditions is establis
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mm) 2524-T3 and 2024-T3 bare sheet
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condition. The corrosion resistance
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gaseous hydrogen) into a digital sy
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chemicals. At one barracks stores s
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20010016874 Leiden Univ., Centre fo
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Department of Defense (DOD) continu
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19990054229 Universidad Carlos 3 de
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lution filter banks, orthonormal an
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20000012199 Defence Evaluation Rese
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tion model of the ZOne DIgital Auto
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also 20000012198; Copyright Waived;
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for tactical data communication bas
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advantage of the anticipated benefi
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20000032373 Electronic Systems Cent
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U.S. Army’s Training and Doctrine
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Marconi Communications S.p.A., Ital
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20000091431 Maribor Univ., Maribor,
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Koehler, Joachim, Gesellschaft fuer
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clutter background. Performances of
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interaction faults. A fault toleran
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Information Management Challenges i
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ative layers and corner singulariti
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added resistance due to waves of Su
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code provides an algorithm for futu
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Kolenikov, A. F., Academy of Scienc
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and vortices. This document present
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19990056423 Honeywell Technology Ce
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targets, multi-angle observations o
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ment. The key elements of the tailo
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ased on radar and optronics. Variou
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20000032661 National Defence Resear
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simulation programmes and wargames
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20010012839 Deutsche Forschungsanst
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Copyright Waived; Avail: CASI; A02,
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discussed. This covers the definiti
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Aeromedical Training including: Sim
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ment or loss of consciousness (G-LO
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attlefield insults and derives esti
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melatonin has probably numerous oth
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operational reasons. There is an un
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peptide, when synthesized locally i
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Requirements; November 2000, pp. 11
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20010032437 Defence Research Agency
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- 17-4; In English; See also 200100
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Vukov, Mirtcho, National Center of
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M., Defence Evaluation Research Age
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data sheds some light on this and a
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20010076843 Air Force Research Lab.
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GID represent some of the commonest
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Diminished funding for research mak
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profiles for altitude simulations a
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time. Hence, counter to introspecti
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present study we conclude that the
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the Military College. During this t
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training performance data were avai
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cause, the integration of selection
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methods are appropriate, and that o
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flying; (4) Specific single task tr
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improve either flight safety and ai
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20000032391 North Dakota State Univ
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tional fluid dynamics simulations a
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phase. We present in this paper a f
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20010056516 Office National d’Etu
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distribution. This ‘alternative
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etween these objective and human-ba
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837-1059-2; Copyright Waived; Avail
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20010066268 North Atlantic Treaty O
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therefore needs to focus on what ha
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19990092814 Thomson-CSF, Radars and
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20000032868 NASA Lewis Research Cen
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discussed in terms of challenges wi
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potential benefits that would arise
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probes the ground until he hits a s
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actual size and he has to mike the
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working groups, conferences, worksh
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aircrew training from Belgium, Fran
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20020016341 Military Univ. of Techn
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ling; vertical accelerations and ve
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locations of sensors and actuators
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social problems, which add to the b
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Germany; Onken, Reiner, Universitae
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forces. The discussion is largely q
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fast and reduces planning time from
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ing discussion areas: (1) Missile f
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20000089820 Wehrtechnische Dienstel
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20010067706 Otto-von-Guericke Univ.
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conforming non-orthogonal structure
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comparable or higher quality by non
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while procedurally based informatio
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gain interoperability between diffe
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address some or all of these diffic
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20020016334 Defence Research Establ
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C4I systems. The article is primari
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ACTIVE CONTROL Subject Term Index A
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AEROSPACE MEDICINE Subject Term Ind
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AIRCRAFT ENGINES Subject Term Index
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ALERTNESS Subject Term Index Sleep
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ATTACK AIRCRAFT Subject Term Index
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C-130 AIRCRAFT Subject Term Index C
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COMBUSTION CHAMBERS Subject Term In
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COMPLEX SYSTEMS Subject Term Index
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COMPUTERIZED SIMULATION Subject Ter
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COSINE SERIES Subject Term Index CO
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DATA PROCESSING Subject Term Index
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DETECTION Subject Term Index Lesson
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EDUCATION Subject Term Index Hypoba
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ENVIRONMENT SIMULATION Subject Term
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FATIGUE (MATERIALS) Subject Term In
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FLIGHT CREWS Subject Term Index Nig
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FUEL COMBUSTION Subject Term Index
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HARMONIC OSCILLATION Subject Term I
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HUMAN FACTORS ENGINEERING Subject T
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IMAGING RADAR Subject Term Index IM
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INTELLIGENCE Subject Term Index INT
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LIQUID FUELS Subject Term Index LIQ
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MATHEMATICAL MODELS Subject Term In
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MILITARY TECHNOLOGY Subject Term In
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NAVY Subject Term Index US Naval an
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OPTIMIZATION Subject Term Index Pro
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PHYSIOLOGICAL EFFECTS Subject Term
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PRESSURE CHAMBERS Subject Term Inde
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QUALITY CONTROL Subject Term Index
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RHYTHM (BIOLOGY) Subject Term Index
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SHOCK TUNNELS Subject Term Index SH
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STANDARDIZATION Subject Term Index
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SYSTEM FAILURES Subject Term Index
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TECHNOLOGY ASSESSMENT Subject Term
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TRAINING SIMULATORS Subject Term In
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VALSALVA EXERCISE Subject Term Inde
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WAVELENGTHS Subject Term Index WAVE
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A Abdi, Frank Rotorcraft Damage Tol
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Personal Author Index BINGEL, P. In
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Personal Author Index CALDWELL, J.
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Personal Author Index CROWLEY, J. S
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Personal Author Index DZIERZANOWSKI
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Personal Author Index FRISE, E. Fel
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Personal Author Index GWILLIAM, N.
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Personal Author Index HUNT, MELVYN
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Personal Author Index KIND, R. J. J
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Personal Author Index LEBLAYE, P. K
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Personal Author Index MARKOU, IOANN
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Personal Author Index MOSS, J. B. M
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Personal Author Index PARK, T. W. N
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Personal Author Index REID, SUSAN A
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Personal Author Index SAWYER, SCOTT
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Personal Author Index THORSEN, ULF
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Personal Author Index VOGT, RYSZARD
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Personal Author Index WYNIA, SIKKE
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AVT - Applied Vehicle Technology Pa
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RTO-TR-038 Ice Accretion Simulation
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ISBN 92-837-1053-3 The Human Factor
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20000020810 - 61 20000020811 - 61 2
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