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This report results from a contract tasking Samara State Aerospace University (SSAU) as follows: The contractor will develop and investigate aerodynamic and structural weight theories associated with morphing wing technology. The contractor has proposed using a generic, telescoping wing design as the baseline configuration for analysis. Specific tasks to be accomplished include: 1. Development of a new weight analysis theory; 2. Creation of an analytical model of the telescoping wing design; 3. Calculation of Load-Carrying Factor for the uniform pressure distribution or for some prescribed spanwise circulation distribution; 4. Calculation of a new dimensionless Load-Carrying Factor'. according to a special algorithm which will allow computing absolute weight and wing weight fracture for different specific wing loadings; 4. Calculation of a Construction Factor' from retrospective analysis of similar wings; 5. Wind tunnel tests of a model of the telescoping wing to assess impact on aerodynamic coefficients and drag polars; 6. Creation of the vortex-lattice computational model of the telescoping wing to obtain its aerodynamic coefficients; 7. Creation of the panel method computational model for detailed study of flow features in the region of the telescope wing Joint; and 8. Analysis and comparison of experimental and computational results., The original document contains color images.

The goal of this Trident Project was the determination of the maximum dynamic lift coefficient of high aspect ratio control surfaces to reduce the uncertainty in design. The effective design of rudders, diving planes and other control surfaces with high aspect ratios, i.e. a larger depth in proportion to chord length, is hindered by uncertainty in the dynamic coefficient of lift of the control surface as it passes through a transient fluid flow affected by the fluid free surface. This unsteady flow creates momentary spikes in loading, which if large enough, can cause immediate or fatigue-induced failure. High aspect ratio control surfaces are more efficient (higher lift-to-drag ratios) than traditional designs and generate greater lift at smaller angles of attack to the incoming flow, however their increasing length produces significantly greater bending moments in the shaft. This has led to a number of in-service failures. As high aspect ratio lifting surfaces become more commonplace on surface vessels and submarines, more precise knowledge of loading conditions is required to avoid more failures. The dynamic lift coefficient is a dimensionless load factor that can be applied to geometrically similar bodies. As the dynamic coefficient of lift is dimensionless, by employing the principles of similitude, the testing of large bodies such as ships hulls or wing sections can be successfully analyzed through the use of scale models tested in wind tunnels or tow tanks. Data collection took place in the David Taylor Model Basin at the Naval Surface Warfare Center. Two 1/5-scale rudders were tested on a donated 15-foot vessel model. Strain gages were attached to the rudder and shaft to determine the strain distribution and overall deformation., The original document contains color images.

This research program has focused on a novel approach to the control of the aerodynamic performance of lifting surfaces by fluidic modification of their apparent aerodynamic shape, or virtual aerosurface shaping, This flow control approach emphasizes fluidic modification of the "apparent" aerodynamic shape of the surface with the objective of altering or prescribing the streamwise pressure gradient. Control is typically effected by the interactions of arrays of synthetic jet actuators with the cross flow that displace the local streamlines near the surface and thereby induces an "apparent" modification of the flow boundary. The operating frequency of the control jets is high enough so that the actuation period at least an order of magnitude lower than the relevant characteristic time scale of the flow. Therefore, the interaction domains between the control jets and the cross flow are quasi-steady and hence the induced aerodynamics forces are virtually time-invariant., The original document contains color images. All DTIC reproductions will be in black and white.

Blade section normal force and pitching moment were investigated for six rotors operating at transition and high speeds: H-34 in flight and wind tunnel, SA 330 (research Puma), SA 349/2, UH-60A full-scale and BO- 105 model (HART-I). The measured data from flight and wind tunnel tests were compared with calculations obtained using the comprehensive analysis CAMRAD II. The calculations were made using two free wake models: rolled-up and multiple-trailer with consolidation models. At transition speed, there is fair to good agreement for the blade section normal force between the test data and analysis for the H-34, research Puma and SA 349/2 with the rolled-up wake. The calculated airloads differ significantly from the measurements for the UH-60A and BO-105. Better correlation is obtained for the UH-60A and BO-105 by using the multiple-trailer with consolidation wake model. In the high speed condition, the analysis shows generally good agreement with the research Puma flight data in both magnitude and phase. However, poor agreement is obtained for the other rotors examined. The analysis shows that the aerodynamic tip design (chord length and quarter chord location) of the Puma has an important influence on the phase correlation., Presented at the American Helicopter Society Decennial Specialist's Conference on Aeromechanics (4th) held in San Francisco, CA on 21-23 January 2004. Published in the Proceedings of the American Helicopter Society Decennial Specialist's Conference on Aeromechanics (4th), 2004.

This paper describes the work achieved during the recent years to improve the quality of Navier-Stokes computations for industrial purpose. The main axes for improvement are the grid adaptation strategy the turbulence modeling and the application to industrial missiles. A few years ago Navier-Stokes computations were rarely run for external aerodynamics simulations. This is no longer true. Both industrial companies (EADS Aerospatiale Matra Missiles and MATRA BAe Dynamics France) now use it on a regular basis. Nevertheless some progresses are still necessary especially for the prediction of the axial force., Also see: ADM001490, Presented at RTO Applied Vehicle Technology Panel (AVT) Symposium held inLeon, Norway on 7-11 May 2001, The original document contains color images.

A computational fluid dynamics (CFD) method has been employed to compute vortical flows around slender wing/body configurations. The emphasis of the paper is on the effectiveness of an adaptive grid procedure in capturing' concentrated vortices generated at sharp edges or flow separation lines of lifting surfaces flying at high angles of attack. The method is based on a tetrahedral unstructured grid technology developed at the NASA Langley Research Center. Two steady-state, subsonic, inviscid and Navier-Stokes flow test cases are presented to demonstrate the applicability of the method for solving practical vortical flow problems. The first test case concerns vortex flow over a simple 650 delta wing with different values of leading-edge radius. Although the geometry is quite simple, it poses a challenging problem for computing vortices originating from blunt leading edges. The second case is that of a more complex fighter configuration. The superiority of the adapted solutions in capturing the vortex flow structure over the conventional unadapted results is demonstrated by comparisons with the wind-tunnel experimental data. The study shows that numerical prediction of vortical flows is highly sensitive to the local grid resolution and that the implementation of grid adaptation is essential when applying CFD methods to such complicated flow problems., Also see: ADM001490, Presented at RTO Applied Vehicle Technology Panel (AVT) Symposium on "Advanced Flow Management: Part A- Vortex Flows and High Angle of Attack for Military Vehicles" held in Loen, Norway on 7-11 May 2001. Pub. in report number RTO- MP-069(I). The original document contains color images.

A simple method of estimating the pitch-damping coefficients for axisymmetric flight bodies has been developed. The method is suitable for preliminary design and parametric studies. The procedure is based on concepts from slender body theory, but the method significantly improves the accuracy of the results through the use of correlation functions. To derive the correlation functions, extensive use of a recently developed and validated computational fluid dynamics (CFD) method for predicting the pitch-damping coefficients has been employed. The CFD method provides important details necessary to derive the correlation functions that are unavailable from the current experimental database. The method has been validated for a variety of nose geometries, body lengths, and Mach numbers using both experimental and computationally generated databases., The original document contains color images. All DTIC reproductions will be in black and white.

Report developed under STTR contract for topic ARMY02-T003. The initial development of a software toolkit to support parallel adaptive numerical analysis is described. The tools are focused on unstructured adaptive techniques which are among the most difficult to implement in parallel. To support dynamic load balancing software was developed to capture and utilize a hardware model, which represents the set of computing resources available for a given problem (analysis) at hand Mesh refinement software was extended for parallel operation, by adding mesh migration steps that enable modifications to be made to mesh entities lying on partition boundaries. The resulting capabilities were demonstrated in a parallel adaptive Navier-Stokes analysis. A new methodology for solution transfer, which combines the update of solution information with the mesh modification, was also implemented and compared to a global solution transfer procedure used in a commercial finite element code. The results showed that the new procedure leads to smaller relative difference in strain norm. Methods for reducing the storage required for adaptively evolving meshes were also investigated. An approach based on storing differences between successive mesh iterations showed that a significant reduction in file space is feasible., The original document contains color images. All DTIC reproductions will be in black and white. Topic no. ARMY02-T003.

Aerodynamic shape design is a challenging application for modem computational fluid dynamic. This is, in-deed, a very productive field, both in term of scientific publications and developed applications. Furthermore it is possible to find excellent analysis/design software released in the public domain. However, industry, in particular automotive and aerospace, but also emerging fields like wind and sea extracted energy, continuously needs increased product performance and competitiveness. This requires increasingly refined, advanced and complex computational models capable of describing flow behavior with higher and higher fidelity. Thus the design tools available to scientist and engineers are challenged by the growing complexity of simulation models. The increasing price to performance ratio of nowadays computers is a strong allied of designers in this though race, but often it is not enough. Objectives of this lecture is to explore and analyze some of the technologies that, together with the increasingly available computational power, may help to exploit the tremendous potential of high fidelity flow field models to aerodynamic shape design challenges. For the sake of simplicity we will concentrate on aerofoil design. Two aspects will be considered: the computer algebra assisted development of code related to gradient computation through adjoint techniques and the use of fitness approximation techniques to improve the performance of search procedures. The concern of this first lecture will be more methodological, while some applications will be illustrated in the next one. The outline of the first part of this lecture is as follows: some short remarks on adjoint problems will introduce the framework in which the Computer Algebra techniques will be mainly used here., See also ADM001519. Published in proceedings, RTO-EN-022, p9-1 through 9-25, 2003. Presented at the RTO AVT Course on Intelligent Systems for Aeronautics held in Rhode-Saint-Genese, Belgium on 13-17 May 2002. The original document contains color images.

This paper was written to provide an overview of the potential of using genetic algorithms (GA's) for aerospace design or to support aerospace design efforts. Genetic algorithms can design individual vehicle components, like the external aerodynamic shape, the internal supporting structure, the propulsion system, the autopilot, or just about any other system component that has requirements that can be mathematically represented and modeled. Genetic algorithms can also be used to support design studies through their ability to find optima for complicated multi-dimensional optimization topologies. This paper, which overviews multiple recent aerospace design efforts using genetic algorithms, was an invited paper presented at a joint NATO/Von-Karmen-Institute Workshop on Intelligent System held in Brussels, Belgium during May of 2002., See also ADM001519. Presented at the RTO AVT Course on "Intelligent Systems for Aeronautics" held in Rhode-Saint-Genese, Belgium on 13-17 May 2002. Published in RTO-EN-022, p8-1-8-20, 2002. The original document contains color images.

The effect of simple lean on an axial turbine stator was examined using a three dimensional viscous flow code. A constant section stator with no meridional contouring was chosen to isolate the effect of the lean. Four different lean angles were examined at three hub/tip ratios.

A quaternion-based attitude control system is developed for the X-33 in the ascent flight phase. A nonlinear control law commands body-axis rotation rates that align the angular velocity vector with an Euler axis defining the axis of rotation that will rotate the body-axis system into a desired-axis system. The magnitudes of the commanded body rates are determined by the magnitude of the rotation error. The commanded body rates form the input to a dynamic inversion-based adaptive/reconfigurable control law. The indirect adaptive control portion of the control law uses online system identification to estimate the current control effectiveness matrix to update a control allocation module. The control allocation nominally operates in a minimum deflection mode; however, if a fault is detected, it can operate in a null-space injection mode that excites and decorrelates the effectors without degrading the vehicle response to enable online system identification. The overall 5 stem is designed to provide fault and damage tolerance for the X-33 on ascent.

Early strike forces of the future military are envisioned as being lightly armored to enable a rapid deployment. The increased vulnerability and operational tempo of lightly armored forces evokes the need for beyond-line-of-sight reconnaissance capability under the control of the troops on the ground. The objective of this work is to explore and evaluate system level concepts that fulfill this mission using an unmanned air/ground vehicle. Extreme Engineering has designed a UAGV according to specifications given by AMCOM. The UAGV is described in detail in this paper. It uses a ducted fan design with a Wankel engine fueled by JPA jet fuel. An aerodynamic shape and advanced fan design allow the vehicle to fly quickly and efficiently. Light, yet strong materials will be used for the vehicle body and wheel system, making this vehicle highly resistant to all types of environment. Communications and sensor equipment will be used to recognize the environment around the vehicle. Three computers will operate on board the UAGV to interpret and respond to the environment. A system has been designed to reduce the vehicle's acoustic signature to a practically undetectable level. Each vehicle system is detailed in this paper, and accompanying calculations, charts, and tables are included., See Also ADA397955 ADA397956 ADA397957 .

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, higher density fuels, rocket motor thrust magnitude control, high thrust motor, and reaction jet control. * Missile airframe materials technologies. Assessments include hypersonic structure materials, composite structure materials, hypersonic insulation materials, multi-spectral domes, and reduced parts count structure., The original document contains color images. Pres: Systems Concepts and I ntegration Panel (SCI) and Consultant and Exchange Programme of RTO, 18-19 Jun 2001, Tbilisi, Georgia(Russia); 21-22 Jun 2001, Bucharest, Romania; 24-26 Jun 2001, Madrid, Spain; 28-29 Jun 2001, Stockholm, Sweden. p2-1/2-19 This article is from ADA394520 Technologies for Future Precision Strike Missile Systems (les Technologies des futurs systemes de missiles pour frappe de precision)

Early strike forces of the future military are envisioned as being lightly armored to enable a rapid deployment. The increased vulnerability and operational tempo of lightly armored forces evokes the need for beyond-line-of-sight reconnaissance capability under the control of the troops on the ground. The objective of this work is to explore and evaluate system level concepts that fulfill this mission using an unmanned air/ground vehicle. Extreme Engineering has designed a UAGV according to specifications given by AMCOM. The UAGV is described in detail in this paper. It uses a ducted fan design with a Wankel engine fueled by JPA jet fuel. An aerodynamic shape and advanced fan design allow the vehicle to fly quickly and efficiently. Light, yet strong materials will be used for the vehicle body and wheel system, making this vehicle highly resistant to all types of environment. Communications and sensor equipment will be used to recognize the environment around the vehicle. Three computers will operate on board the UAGV to interpret and respond to the environment. A system has been designed to reduce the vehicle's acoustic signature to a practically undetectable level. Each vehicle system is detailed in this paper, and accompanying calculations, charts, and tables are included., See Also ADA397955 ADA397956 ADA397958 .

Early strike forces of the future military are envisioned as being lightly armored to enable a rapid deployment. The increased vulnerability and operational tempo of lightly armored forces evokes the need for beyond-line-of-sight reconnaissance capability under the control of the troops on the ground. The objective of this work is to explore and evaluate system level concepts that fulfill this mission using an unmanned air/ground vehicle. Extreme Engineering has designed a UAGV according to specifications given by AMCOM. The UAGV is described in detail in this paper. It uses a ducted fan design with a Wankel engine fueled by JPA jet fuel. An aerodynamic shape and advanced fan design allow the vehicle to fly quickly and efficiently. Light, yet strong materials will be used for the vehicle body and wheel system, making this vehicle highly resistant to all types of environment. Communications and sensor equipment will be used to recognize the environment around the vehicle. Three computers will operate on board the UAGV to interpret and respond to the environment. A system has been designed to reduce the vehicle's acoustic signature to a practically undetectable level. Each vehicle system is detailed in this paper, and accompanying calculations, charts, and tables are included., See Also ADA397955 ADA397957 ADA397958 .

Early strike forces of the future military are envisioned as being lightly armored to enable a rapid deployment. The increased vulnerability and operational tempo of lightly armored forces evokes the need for beyond-line-of-sight reconnaissance capability under the control of the troops on the ground. The objective of this work is to explore and evaluate system level concepts that fulfill this mission using an unmanned air/ground vehicle. Extreme Engineering has designed a UAGV according to specifications given by AMCOM. The UAGV is described in detail in this paper. It uses a ducted fan design with a Wankel engine fueled by JPA jet fuel. An aerodynamic shape and advanced fan design allow the vehicle to fly quickly and efficiently. Light, yet strong materials will be used for the vehicle body and wheel system, making this vehicle highly resistant to all types of environment. Communications and sensor equipment will be used to recognize the environment around the vehicle. Three computers will operate on board the UAGV to interpret and respond to the environment. A system has been designed to reduce the vehicle's acoustic signature to a practically undetectable level. Each vehicle system is detailed in this paper, and accompanying calculations, charts, and tables are included., See Also ADA397956 ADA397957 ADA397958 .

Our long-term objective has been the development of a high-fidelity and high-performance simulation capability for predicting and optimizing the dynamic aeroelastic response of a fighter during three-dimensional high-G maneuvers in subsonic, transonic, and supersonic airstreams. Our focus has been on Air Force problems involving a modern fighter or bomber, and relevant to new approaches for flutter testing, mitigation of limit-cycle (LCO) and pilot induced (P1O) oscillations, as well as performance optimization. Our starting point has been the unique aeroelastic simulation capability developed at the University of Colorado under the sponsorship of the Air Force office of Scientific Research, and in partnership with the Flight Test Center at the Edwards Air Force Base.

An Improved Semi empirical Method has been developed to estimate the static aerodynamics of configurations that use a trailing edge flap for control. The method is based on deflecting the full aft-located lifting surface an amount that allows the normal force coefficient to be equal to that generated by the deflected flap. A transfer in pitching moments and a modified axial force coefficient is derived to complete the set of static aerodynamics. The method is derived using theoretical methods that are a part of the 1998 version of the Naval Surface Warfare Center aeroprediction code and two sets of experimental data. Comparison of the improved method to available data shows the method to give satisfactory results over the practical range that trailing edge flaps are contemplated for use. However, additional wind tunnel data is needed to refine and expand the applicability of the method. This is particularly true for transonic Mach numbers and for supersonic Mach numbers where the angle of attack and control deflection are of the same sign.

The U.S. Army Research Laboratory (ARL) is interested in applying state-of-the-art computational tools to the study of projectile aerodynamics at angle of attack and moderate Mach number. The WIND flow solver has been used to study the aerodynamics of two missile configurations. Computations have been done for several turbulence models. Numerical results are compared with experimental data for body surface pressures, flow field pitot pressures, and body loads., Original contains color plates: All DTIC reproductions will be in black and white. Prepared in cooperation with the US Military Academy, West Point, NY 10996.

This study evaluated the role of adaptive cueing interfaces as a means of integrating visual and auditory displays for target designation and was conducted at the Air Force Research Laboratory's Synthesized Immersion Research Environment (SIRE) facility. Twelve pilots with a mean of 2652 flight hours performed a simulated flight task in which they were instructed to maintain a prescribed flight path, air speed, and altitude. Pilots visually searched out-the-window for ground and air targets. Seven target-location cueing conditions were employed which featured spatialized auditory, visual, non-adaptive multimodal, and adaptive multimodal display configurations. Designation accuracy was poorest in the non-cueing and spatialized auditory display conditions in comparison to the other display configurations. The addition of spatialized sound to visual cueing reduced target designation time in comparison to the visual cue alone in some conditions, and the time advantage form multimodal cueing was approximately 825 msec. Multimodal cueing also reduced head motion and lowered pilots' workload by approximately 30%. This study suggests that multimodal cueing may effectively aid target localization in tactical aviation and revealed no advantage or disadvantage for presenting multimodal information adaptively over presenting it in a fixed format; the benefits associated with multimodal information were identical in both fixed and adaptive formats.

Practical experience with airfoil optimization techniques has revealed unexpected difficulties. Traditionally the performance of an airfoil is optimized for given, or assumed, model parameters. Experience has indicated that a deterministic optimization for discrete operating conditions may result in dramatically inferior performance when the actual conditions are difficulties different from these, somewhat arbitrarily chosen, design values. Extensions to multi-point optimization have proven unable to adequately remedy the problem of "localized optimization". This paper presents an intrinsically statistical approach and demonstrates how the shortcomings of multi-point optimization with respect to "localized optimization" can be overcome., Prepared in cooperation with The College of William and Mary, Williamsburg, VA.

The main objective of this has been the development awl maturation of a adjoint-based, viscous design techniques for Aerodynamic Shape Optimization (ASO) of complete aircraft configurations. We have followed a systematic approach in the development, validation and testing of our methods. In addition, we have carried out some preliminary work in the development of a mathematical environment for high-fidelity aero-structural optimization as a first step towards the realization of a high-fidelity multidisciplinary optimization disability. A number of significant milestones achieved in this process are detailed in the final report.

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. * Missile airframe materials technologies. Assessments include hypersonic structure materials, composite structure materials, hypersonic insulation materials, multi-spectral domes, and reduced parts count structure., Presented at the Systems Concepts and Integration Panel (SCI) and the Consultant and Exchange Programme of RTO in support of a Lecture Series presented on 23-24 March 2000 in Atlanta, USA, on 3-4 April 2000 in Turin, Italy, and on 6-7 April 2000 in Ankara, Turkey. This article is from ADA387602 Technologies for Future Precision Strike Missile Systems (les Technologies des futurs systemes de missiles pour frappe de precision)

The F-5 test series (see chapter 5) provides a succession of geometries of increasing complexity Ref. 1 Ref. 2, which will he useful for validating CFD codes during their development. In this chapter a range of CFD results are provided for the clean wing configuration at selected flow conditions, and a more limited set for one complex configuration. Results from essentially slate of the art UTSP (Unsteady Transonic Small Perturbation), Full Potential, Euler. and NS (Navier-Stokes) codes are presented. this will allow the reader to gauge anticipated modelling accuracy for code development purposes. Table 1 summarises the methods used by contributors reported herein, the methods themselves are described in a standard proforma and the results collated as a series of plots. The flow conditions calculated are summarised in Table 2 and Table 4. Two or more methods are presented for each level of modelling approxmation in order to assist the reader in gauging the likely level of variation in solution at a particular level of approximation., The original document contains color images. This article is from ADA390566 Verification and Validation Data for Computational Unsteady Aerodynamics (Donnees de verification et de valadation pour l'aerodynamique instationnaire numerique)

This data set relates to tests at RAE which were carried out and reported by D (3 Mabey, B L Welsh and B E Cripps, ref. 1. The tests were undertaken to provide data for the validation of codes for the prediction of both steady and unsteady pressures on low aspect ratio configurations, suitable for the wings or controls of military aircraft. Comprehensive measurements have not been available to verify such codes, although some measurements were obtained during the NORA programmer. This was a collaborative test on a low aspect ratio model oscillating about a swept axis, with the main aim of investigating dynamic interference in transonic wind tunnels. NORA was named after the participating organizations: NLR, ONERA, RAE, and AVA (branch of DFVLR). For the verification of transonic theories, a serious limitation of the NORA tests was that the steady and unsteady pressures were measured at different sections, with only a few measurements at each section. To overcome the lack of comprehensive measurements on a low aspect ratio configuration it was decided to make extensive measurements of steady and unsteady pressures on a model of the AGARD SMP (Structures and Materials Panel) tailplane, which has a planform similar to that of the wings and controls used on many military aircraft. Previous tests have shown that for experiments in time-dependent aerodynamics it is essential to minimise aeroelastic distortion when the model is driven. To avoid measured pressures with a significant contribution due to the distortion in the present tests, the model had to move almost as a rigid body when it was oscillated at high frequencies. Hence the model was constructed in carbon fibre, which provided both high stiffness and low inertia. The high stiffness was aided by the 1 0% thickness of the section used, which is significantly thicker than the sections usually employed on combat aircraft., The original document contains color images, p271-294. This article is from ADA390566 Verification and Validation Data for Computational Unsteady Aerodynamics (Donnees de verification et de valadation pour l'aerodynamique instationnaire numerique)

Decomposition of a flow domain into blocks, so-called domain decomposition, is the most challenging task in the process of multi-block grid generation. In this paper, a method is described to facilitate the construction of the blocks in exterior flow domains. Layers of blocks are automatically generated by solving the Laplace equation in a region between the configuration surface and a bounding box surface. A boundary element method based on simple-source distributions is used to solve the Laplace equation. After solving the Laplace equation, starting with a face decomposition of the configuration surface, blocks are generated by marching along the electrostatic field lines until a user specified potential is reached. The process may be repeated to generate a second layer of blocks. However, for configurations with a strongly non-convex shape, some blocks may need to be added manually to prevent deterioration of the block shapes in the second-layer. Therefore, the method is called semi-automatic., Text in English. Also available for download at the NLR homepage http://www.nir.nl/public/library/index.html. Presented at the 7th International Conference on Numerical Grid Generation in Computational Field Simulations, Whistler, Canada, 25-28 Sep 2000.

This paper addresses the problem to define a methodology for the analysis of the performances of different aircraft configurations in the phase of conceptual design. The proposed approach is based on a numerical optimization procedure where a scalar objective junction, the take-off weight, is minimized. The optimization algorithm has obviously important consequences both from the point of view of the computational times and of the obtained results. For this reason a preliminary discussion is made where various different methodologies are critically compared. Although the best compromise between different approaches is probably given by an integration between a genetic algorithm approach and a classical gradient method, in this phase only the latter procedure has been used to perform the simulations. The methodology takes into account the high number of geometrical parameters and the flight mechanics, requirements involved in the problem. A basic example is described, and the use of the proposed methodology to investigate the effects of different geometrical and technological parameters is discussed., Presented at Symposium of RTO Applied Vehicle Technology Panel (AVT) Canada, 18-21 Oct 1999, p6-1/6-10. This article is from ADA388284 Aerodynamic Design and Optimisation of Flight Vehicles in a Concurrent Multi-Disciplinary Environment (la Conception et l'optimisation aerodynamiques des vehicules aeriens dans un environnement pluridisciplinaire et simultane)

The Symposium dealt with Design Issues and more specifically Aerodynamic Design and Optimization of Flight Vehicles in a Concurrent Multi-Disciplinary Environment. Thirty two papers and a Keynote address were presented with the objective to survey the current and future scene given the trend towards a more concurrent and multi-disciplinary approach to aerospace vehicle engineering. There were six sessions covering the following topics: * Lessons Learnt/Requirements for the Future Regard to the Role of Aerodynamicists in a Concurrent Multi-Disciplinary Design Process * The Role of Aerodynamics in Concept Phase of a Project Design * MDO and the Aerodynamics Design Process * Methodologies/Tools for Aerodynamic Optimization * Application of Methodologies/Tools for Aerodynamic Optimization Techniques for Rapid Database Generation, Presented at Symposium of RTO Applied Vehicle Technology Panel (AVT) Canada, 18-21 Oct 1999. For individual articles, see ADP010499 thru ADP010530

A series of flight tests were conducted on the PC 9/A aircraft, A23-045, at the Royal Australian Air Force's Aircraft Research and Development Unit. System identification techniques were applied to the data obtained from these flight tests to determine the stability and control derivatives of the aircraft. The longitudinal and lateral results for the aircraft in approach and departure configurations are presented in this report and comparisons are made with wind tunnel estimates., Original contains color plates: All DTIC reproductions will be in black and white.

A numerical methodology for the evaluation of aerodynamic loads acting on a complex lifting configuration is presented. The work is limited to the case of attached high-Reynolds number flows. A viscous/potential interaction technique is utilized to take into account the effects of the viscosity. For the potential-flow analysis, a boundary element formulation is used; for simplicity, only incompressible flows are examined. The theoretical basis of the present methodology is briefly described. Comparisons with available, numerical and experimental results are included., Presented at Symposium of RTO Applied Vehicle Technology Panel (AVT) Canada, 18-21 Oct 1999, p16-1/16-11. This article is from ADA388284 Aerodynamic Design and Optimisation of Flight Vehicles in a Concurrent Multi-Disciplinary Environment (la Conception et l'optimisation aerodynamiques des vehicules aeriens dans un environnement pluridisciplinaire et simultane)

With the capacity of today's computers; one can envisage the resolution of shape optimization problems in aerodynamics. Nevertheless, optimization methods require many evaluations of different aerodynamic configurations, and so are much more expensive than a single analysis. It is therefore mandatory to find methods that evaluate aerodynamic functions and their gradient at the lowest possible computational cost; as well as fast and robust optimization methods., Presented at Symposium of RTO Applied Vehicle Technology Panel (AVT) Canada, 18-21 Oct 1999, p18-1/18-12. This article is from ADA388284 Aerodynamic Design and Optimisation of Flight Vehicles in a Concurrent Multi-Disciplinary Environment (la Conception et l'optimisation aerodynamiques des vehicules aeriens dans un environnement pluridisciplinaire et simultane)

The purpose of this paper is to report on methods which have been developed or which are under development at Alenia Aeronautica for multidisciplinary optimum de- sign, with particularly emphasis on aerodynamic shape design. Results of transonic 2D and 3D optimization problems are presented., Presented at Symposium of RTO Applied Vehicle Technology Panel (AVT) Canada, 18-21 Oct 1999, 20-1/20-7. This article is from ADA388284 Aerodynamic Design and Optimisation of Flight Vehicles in a Concurrent Multi-Disciplinary Environment (la Conception et l'optimisation aerodynamiques des vehicules aeriens dans un environnement pluridisciplinaire et simultane)

This paper addresses applied procedures for nonlinear aerodynamic model development and extraction from flight data for the S-3B Viking aircraft. The entire analysis procedures, from dynamic flight test data management to final blending and validation of the upgraded aerodynamic model, was performed within the Integrated Data Evaluation and Analysis Systems (IDEAS) developed by SAIC. IDEAS is a powerful database management system and analysis software containing a full complement of flight data preprocessing, calibration, simulation, model estimation, model verification, and validation tools.

Since early in the development of aerodynamic computational tools, when complex aircraft configurations could for the first time be accurately represented, the Holy Grail of the CFD community became predicting the trajectory of a store from an aircraft. However, the participants could never agree on either what would constitute a successful prediction, or what computational approach should be used. Many claimed that store separation was an unsteady phenomenon, and that the solutions would have to be done in a time dependent manner (this despite the multitude of evidence to the Contrary, where quasi-steady wind tunnel simulations in almost all cases have shown excellent agreement with the flight test results).

A series of flight tests were conducted on the PC 9/A aircraft, A23-045 at the Royal Australian Air Force's Aircraft Research and Development Unit System identification techniques were applied to the data obtained from these flight tests to determine the stability and control derivatives of the aircraft. The longitudinal results for the aircraft in cruise configuration are presented in this report and comparisons are made with empirical and ground based estimates.

New capability has been added to the NSWC aeroprediction code to allow aerodynamics to be predicted for Mach numbers up to 20 for configurations with flares. This new capability includes extending the static aerodynamic predictions for Mach numbers less than 1.2, improving the body alone pitch damping for Mach numbers above 2.0, and developing a new capability for pitch damping of flared configurations at Mach numbers up to 20. This new capability for flared configurations was validated for several different configurations in the Mach number range of 2 to 8.8. In general, pitch damping predictions of the improved capability was within 20 percent of either experimental data or computational fluid dynamics calculations. This accuracy level is believed to be quite adequate for dynamic derivatives in the preliminary design stage. These new additions to the aeroprediction code will be transitioned to users as part of the 2002 version of the code (APO2).

Our long term objective is the development of a fast and sound methodology for simulating the dynamic aeroelastic response of an aircraft during three-dimensional high-G maneuvers in transonic and supersonic airstreams. Our focus is on AIR FORCE relevant problems exemplified by a transonic/supersonic lifting body geometry performing complex multi-G maneuvers at a free-stream Mach number M00 ranging between 0.7 and 3 (0.7

In a continuation of ongoing Naval Postgraduate School efforts to study the performance characteristics of waverider configured vehicles, a computational fluid dynamic (CFD) analysis of the Mach 6 L/DI(sp)(max) optimized Price waverider was conducted. This analysis was performed to determine the theoretical force and moment data over a broad Mach number spectrum, and to compare theoretical and experimental results in the subsonic flight regime. The CFD determination of force and moment data represents a continuation of the ongoing analysis of the Price waverider configuration. Selected viscous and inviscid flow solutions for flight conditions in the range 0.3 < M(infinity) <6.0, as well as a subsonic (M(infinity) = 03) angle of attack sweep, were conducted using NASA CFD software (OVERFLOW1.8b). Examination of the computed converged flow field solutions suggests that the surface pressure distributions and Mach number contours surrounding the holy are valid. Low speed force and moment coefficient data are shown to exhibit reasonable agreement with the available subsonic wind tunnel data. Additionally, supersonic CFD results show the development of the expected shock layer, exhibiting an attached shock bed at the design Mach number (M(infinity) = 6.0). Evaluation of the computed Mach number effects on lift and drag coefficients at subsonic, transonic and supersonic Mach numbers suggests that the Price waverider may exhibit some flight instabilities across the flight Mach number spectrum.

A modular flight control system is developed for a tailless fighter aircraft with innovative control effectors. Dynamic inversion control synthesis is used to develop a full envelope flight control law. Minor dynamic inversion command variable revisions are required due to the tailless nature of the configuration studied to achieve nominal stability and performance. Structured singular value and simulation analysis shows that robust stability is achieved and robust performance is slightly deficient due to modeling errors. A multi-branch linear programming based method is developed and used for allocation of redundant limited control effectors.

Results of investigations of combustion of liquid hydrocarbon fuel jets aerated by air in a supersonic flow in the two-dimensional diverging-area supersonic combustor are presented. Direct-connect combustor tests were conducted at the combustor entrance Mach number M=2.5 and total temperature in the range of Tt=1635-1742K. The liquid hydrocarbon fuel supply into the combustor was executed through the modified (in comparison with the baseline ) aeroramp configurations:1) installation of baseline aeroramps at the upper and bottom walls of the combustor, 2) combined fuel feed through injector nozzles of aeroramps separately of air and of fuel aerated by air, 3) injectors with the different from baseline aeroramps angles of fuel injection and geometry nozzles. The results obtained with the modified aeroramp configurations are superior to earlier ones obtained with baseline aeroramp configuration.

Any external aircraft modification has potentially far-reaching effects on the capability of the aircraft to succeed or fail in its mission. The authors take a systematic look at the effects that small changes can have upon the whole, with a series of examples that demonstrate why careful review of data or testing is often vital in the assessment of system modifications.

This work discusses an approach, the Approximation Management Framework (AMF), for solving optimization problems that involve computationally expensive simulations. AMF aims to maximize the use of lower fidelity, cheaper models in iterative procedures with occasional, but systematic, recourse to higher fidelity, more expensive models for monitoring the progress of the algorithm. The method is globally convergent to a solution of the original, high fidelity problem. Three versions of AMF, based on three nonlinear programming algorithms, are demonstrated on a 3D aerodynamic wing optimization problem and a 2D airfoil optimization problem. In both cases Euler analysis solved on meshes of various refinement provides a suite of variable fidelity models. Preliminary results indicate threefold savings in terms of high fidelity analyses in case of the 3D problem and twofold savings for the 2D problem.

This report is a part of the documentations which describe the complete development of an STTR Phase 2 effort entitled, Development of the Aerodynamic/Aeroservoelastic Modules in ASTROS. This report is one of four manuals that comprise the final report. The remaining reports consist of the ZAERO User's Manual (Volume 1), the ZAERO Programmer's Manual (Volume 3) and the ZAERO Theoretical Manual (Volume 4). ASTROS is the seamless integration of the ZAERO module into ASTROS. As an aerodynamic enhancement to ASTROS, ZAERO is the ZONA aerodynamic module, unified for all Mach number ranges. This manual assumes the reader is familiar with the ASTROS system (Version 10.0), its terminology and user interface. It provides the complete user interface to the ASTROS system required for preparation of input data. First, an overview of the ZAERO software is presented. Second, the executive control options are described. Third, a complete and comprehensive description of the bulk data input is presented. Fourth, ZAERO modeling guidelines are provided with detailed descriptions for proper aerodynamic, geometry and spline input. Finally, the output description of ZAERO is presented with output samples., ADA370434 ADA370435 ADA370436

This report is a part of the documentations which describe the complete development of an STTR Phase 2 effort entitled, Development of the Aerodynamic/Aeroservoelastic Modules in ASTROS. This report is one of four manuals that comprise the final report. The remaining reports consist of the ZAERO User's Manual (Volume 1), the ZAERO Programmer's Manual (Volume 2) and the ZAERO Applications Manual (Volume 3). ASTROS is the seamless integration of the ZAERO module into ASTROS. As an aerodynamic enhancement to ASTROS, ZAERO is the ZONA aerodynamic module, unified for all Mach number ranges. This manual assumes the reader is familiar with the ASTROS system (Version 10.0), its terminology and user interface. It provides the complete user interface to the ASTROS system required for preparation of input data. First, an overview of the ZAERO software is presented. Second, the executive control options are described. Third, a complete and comprehensive description of the bulk data input is presented. Fourth, ZAERO modeling guidelines are provided with detailed descriptions for proper aerodynamic, geometry and spline input. Finally, the output description of ZAERO is presented with output samples., ADA370433 ADA370434 ADA370435

This report is a part of the documentations which describe the complete development of an STTR Phase 2 effort entitled, Development of the Aerodynamic/Aeroservoelastic Modules in ASTROS. This report is one of four manuals that comprise the final report. The remaining reports consist of the ZAERO Programmer's Manual (Volume 2), the ZAERO Applications Manual (Volume 3) and the ZAERO Theoretical Manual (Volume 4). ASTROS is the seamless integration of the ZAERO module into ASTROS. As an aerodynamic enhancement to ASTROS, ZAERO is the ZONA aerodynamic module, unified for all Mach number ranges. This manual assumes the reader is familiar with the ASTROS system (Version 10.0), its terminology and user interface. It provides the complete user interface to the ASTROS system required for preparation of input data. First, an overview of the ZAERO software is presented. Second, the executive control options are described. Third, a complete and comprehensive description of the bulk data input is presented. Fourth, ZAERO modeling guidelines are provided with detailed descriptions for proper aerodynamic, geometry and spline input. Finally, the output description of ZAERO is presented with output samples., ADA370433 ADA370435 ADA370436

This report is a part of the documentations which describe the complete development of an STTR Phase 2 effort entitled, Development of the Aerodynamic/Aeroservoelastic Modules in ASTROS. This report is one of four manuals that comprise the final report. The remaining reports consist of the ZAERO User's Manual (Volume 1), the ZAERO Applications Manual (Volume 3) and the ZAERO Theoretical Manual (Volume 4). ASTROS is the seamless integration of the ZAERO module into ASTROS. As an aerodynamic enhancement to ASTROS, ZAERO is the ZONA aerodynamic module, unified for all Mach number ranges. This manual assumes the reader is familiar with the ASTROS system (Version 10.0), its terminology and user interface. It provides the complete user interface to the ASTROS system required for preparation of input data. First, an overview of the ZAERO software is presented. Second, the executive control options are described. Third, a complete and comprehensive description of the bulk data input is presented. Fourth, ZAERO modeling guidelines are provided with detailed descriptions for proper aerodynamic, geometry and spline input. Finally, the output description of ZAERO is presented with output samples., ADA370433 ADA370434 ADA370436

The U.S. Army Research Laboratory, NASA Glenn Research Center, and Rolls-Royce Allison are working collaboratively to demonstrate the benefits and viability of a wave-rotor-topped gas turbine engine. The self-cooled wave rotor is predicted to increase the engine overall pressure ratio and peak temperature by 300% and 25 to 30%, respectively, providing substantial improvements in engine efficiency and specific power. Such performance improvements would significantly reduce engine emissions and the fuel logistics trails of armed forces. Progress towards a planned demonstration of a wave-rotor-topped Rolls-Royce Allison model 250 engine has included completion of the preliminary design and layout of the engine, the aerodynamic design of the wave rotor component and prediction of its aerodynamic performance characteristics in on- and off-design operation and during transients, and the aerodynamic design of transition ducts between the wave rotor and the high pressure turbine. The topping cycle increases the burner entry temperature and poses a design challenge to be met in the development of the demonstrator engine., Prepared in collaboration with Dynacs Engineering Company, Inc., Brook Park, Ohio and Roll-Royce Allison, Indianapolis, Indiana.

In order to effectively port production codes originally written for vector processors to reduced instruction set (RISC) based parallel computers, different paradigms have been tried by the parallel computing community. Among the techniques used are message passing and loop level parallelization using hand inserted compiler directives or automatic parallelizing compiler flags. The goals of this report are: (1) to investigate the performance of message passing and loop level parallelization techniques, as they were implemented in the computational fluid dynamics (CFD) code overflow, and (2) to validate the sequential and parallel results obtained on a demonstration problem of interest to the Army that is, a generic missile body configuration. The computational simulations were run, and performance data were gathered on a Silicon Graphics Incorporated (SGI) Power Challenge Array (PCA) and Origin 2000.