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A status report is provided on the collaboration between the Royal Institute of Technology in Sweden and the NASA Langley Research Center regarding the aeroelastic analyses of a full-span fighter configuration wind-tunnel model. This wind-tunnel model was tested in the Transonic Dynamics Tunnel in the summer of 2016. Large amounts of data were acquired including steady/unsteady pressures, accelerations, strains, and measured dynamic deformations. The aeroelastic analyses presented include linear aeroelastic analyses, CFD steady analyses, and analyses using CFD-based reduced-order models. The reduced-order model results also include a comparison of the aeroelastic response of the model in free air and in a computational mesh of the Transonic Dynamics Tunnel in order to determine, computationally, the effects of the wind tunnel on the aeroelastic response., QC 20181221

A status report is provided on the collaboration between the Royal Institute of Tech- nology (KTH) in Sweden and the NASA Langley Research Center regarding the aeroelastic analyses of a full-span fighter configuration wind-tunnel model. This wind-tunnel model was tested in the Transonic Dynamics Tunnel (TDT) in the summer of 2016. Large amounts of data were acquired including steady/unsteady pressures, accelerations, strains, and measured dynamic deformations. The aeroelastic analyses presented include linear aeroelastic analyses, CFD steady analyses, and analyses using CFD-based reduced-order models (ROMs)., QC 20220301

The goal of this project is to integrate novel numerical modeling and computer hardware approaches to compute the non-linear aerodynamic coupling between the ship and aircraft in such a way that execution times are at real-time speeds, allowing for pilot-in-the-loop CFD to be integrated in the piloted flight simulation environment. To achieve the speed gains required, three areas will be targeted for implementation into the CFD simulation framework: (1) numerical algorithms, (2) novel domain boundaries, and (3) Graphical Processing Unit (GPU) hardware. A framework will be established to link the CFD with real time simulations. A building block approach will be employed to first demonstrate non-real time integration of the CFD simulation framework with helicopter flight dynamic models, then real time execution for a minimum fidelity airwake/aircraft simulation, then build to higher fidelity real time simulations. During this reporting period, solver development was performed in two areas: (1) the incompressible formulation employed by the CRUNCH CFD code for CFD airwake simulations was implemented in the structured CRAFT CFD solver (which previously only included a fully compressible formulation) to improve the structured airwake solutions, and (2) an approach has been developed for integrating both the CRAFT CFD and CRUNCH CFD solvers within a common simulation framework (this will make possible combining solver speedups from tasks 1 to 4 in the future)., ONR #BAA12-SN-0028: 2012 Basic and Applied Research in Sea-Based Aviation. The original document contains color images.

The goal of this project is to integrate novel numerical modeling and computer hardware approaches to compute the non-linear aerodynamic coupling between the ship and aircraft in such a way that execution times are at real-time speeds, allowing for pilot-in-the-loop CFD to be integrated in the piloted flight simulation environment. To achieve the speed gains required, three areas will be targeted for implementation into the CFD simulation framework: (1) numerical algorithms, (2) novel domain boundaries, and (3) Graphical Processing Unit (GPU) hardware. A framework will be established to link the CFD with real time simulations. A building block approach will be employed to first demonstrate non-real time integration of the CFD simulation framework with helicopter flight dynamic models, then real time execution for a minimum fidelity airwake/aircraft simulation, then build to higher fidelity real time simulations. During this reporting period, structured simulations for the example LHD airwake were initiated using the improved incompressible formulation implemented in the CRAFT CFD solver that was presented in the previous report. The goal of these simulations is to provide a verification of the structured airwake solution against standard practice unstructured results for the ship platform of interest: LHD., ONR #BAA12-SN-0028: 2012 Basic and Applied Research in Sea-Based Aviation. The original document contains color images.

This project is performed under the Office of Naval Research program on Basic and Applied Research in Sea- Based Aviation (ONR BAA12-SN-0028). This project addresses the Sea Based Aviation (SBA) virtual dynamic interface (VDI) research topic area Fast, high-fidelity physics-based simulation of coupled aerodynamics of moving ship and maneuvering rotorcraft . The work is a collaborative effort between Penn State, NAVAIR, and Combustion Research and Flow Technology (CRAFT Tech). This document presents progress at Penn State University. All software supporting piloted simulations must run at real time speeds or faster. This requirement drives the number of equations that can be solved and in turn the fidelity of supporting physics based models. For real-time aircraft simulations, all aerodynamic related information for both the aircraft and the environment are incorporated into the simulation by way of lookup tables. This approach decouples the aerodynamics of the aircraft from the rest of its external environment. For example, ship airwake are calculated using CFD solutions without the presence of the helicopter main rotor. The gusts from the turbulent ship airwake are then re-played into the aircraft aerodynamic model via look-up tables. For up and away simulations, this approach works well. However, when an aircraft is flying very close to another body (i.e. a ship superstructure), aerodynamic coupling can exist. The main rotor of the helicopter distorts the flow around the ship possibly resulting significant differences in the disturbance on the helicopter. In such cases it is necessary to perform simultaneous calculations of both the Navier-Stokes equations and the aircraft equations of motion in order to achieve a high level of fidelity. This project will explore novel numerical modeling and computer hardware approaches with the goal of real time, fully coupled CFD for virtual dynamic interface modeling & simulation.

In the eld of three-dimensional modeling, we continually struggle to quantify how closely the resulting model matches the physical object being represented. When precision measurements are required, they are often left to high-end, industrial systems. The aim of this thesis is to quantify the level of precision that can be obtained from commodity systems such as the Microsoft Kinect paired with the KinectFusion algorithm. Although the Kinect alone is considered a noisy sensor, the KinectFusion algorithm has shown the ability to build detailed surface models through the aggregation of depth information taken from multiple perspectives. This work represents the first rigorous validation of the three- dimensional modeling capabilities of the KinectFusion algorithm. One experiment is performed to measure the effects of key algorithm parameters such as resolution and range, while another is performed to measure the lower bounds at which objects can be detected and accurately modeled. The first experiment found that the KinectFusion algorithm reduced the uncertainty of the Kinect sensor alone from 10 mm to just 1.8 mm. Furthermore, the results of the second experiment demonstrate that the KinectFusion algorithm can detect surface deviations as little as 1.3 mm, but cannot accurately measure the deviation. Such results form an initial quantification of the KinectFusion algorithm, thus providing confidence about when and when not to utilize the KinectFusion algorithm for precision modeling. The hope is that this work will open the door for the algorithm to be used in real-world applications, such as alleviating the tedious visual surface inspections required for USAF aircraft., The original document contains color images.

We present a time-resolved, adaptive finite element method for aerodynamics, together with the results from the HiLiftPW-2 workshop, where this method is used to compute the flow past a DLR-F11 aircraft model at realistic Reynolds number. The mesh is automatically constructed by the method as part of the computation, and no explicit turbulence model is needed. The effect of unresolved turbulent boundary layers is modeled by a simple parametrization of the wall shear stress in terms of the skin friction. In the extreme case of very high Reynolds numbers we approximate the small skin friction by zero skin friction, corresponding to a free slip boundary condition, which results in a computational model without any model parameter that needs tuning. Thus, the simulation methodology by- passes the main challenges posed by high Reynolds number CFD: the design of an optimal computational mesh, turbulence (or subgrid) modeling, and the cost of boundary layer res- olution. The results from HiLiftPW-2 presented in this report show good agreement with experimental data for a range of different angles of attack, while using orders of magnitude fewer degrees of freedom than what is needed in state of the art methods such as RANS., Part of proceedings: ISBN 978-1-62410-256-1Duplicate record in Scopus 2-s2.0-84902821079QC 20211115

Assessing prediction confidence and enabling its use as a decision-making metric for autonomous model fidelity selection is essential to the USAF's vision of a Digital Twin as a viable approach for condition-based fleet management by tail number. Significant strides have been made in modeling complex interactions of the multi-physics, fluid-thermal-structural coupling applicable to hypersonic flow conditions. However, validation of these models remains a challenge due to limited experimental data for hypersonic conditions. This research addresses quantifying errors and assessing the confidence in aerodynamic pressure and heating predictions for a spherical dome protruding from a flat ramp. Well-characterized aerothermal test data from hypersonic wind tunnel experiments are used to calibrate uncertain model parameters and quantify errors through Bayesian techniques. A Bayesian hypothesis testing-based confidence metric is employed to compare the accuracy in various model predictions. A model selection study is performed for 1st-, 2nd-, and 3rd-order piston theories. The results showed that the greatest confidence in model predictions does not necessarily correspond to the highest-order model., Prepared in collaboration with Vanderbilt University, Nashville, TN. The final version of this conference paper was published in the Proceedings of the AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference (53rd) , held in Honolulu, HI from April 23 through April 26, 2012. The original document contains color images.

This report gives details of an experimental research investigation carried out in the DSTO water tunnel to see whether it is feasible to measure meaningful aircraft static and dynamic aerodynamic derivatives. These derivatives represent the aerodynamic damping and coupling forces and moments on an aircraft and are used in its equations of motion. A Standard Dynamics Model (SDM), a simplified fighter aircraft configuration, was used for the tests. The SDM was subjected to forced small (0.5) sinusoidal pitching oscillations and derivatives were computed from measured model loads, angles of attack, reduced frequency of oscillation and aircraft geometrical parameters. The derivatives obtained in the water tunnel were compared with corresponding published data obtained using SDMs in wind tunnels. Although wind tunnels are the preferred option to obtain derivatives, it was found that it is feasible to use a water tunnel to obtain approximate derivatives, at least for models having SDM-type geometries, especially if derivatives are required quickly and cheaply.

Water tunnels are emerging as a possible useful alternative to small low-speed wind tunnels for an expanded range of aerodynamic testing. In this report, an assessment is made regarding the extent to which water tunnels can be used for such testing. It was found that their suitability for testing given models needs to be assessed on a case-by-case basis. For conventional tests on aircraft, such as force and moment measurements, they compare unfavourably with similar-sized wind tunnels, due to a mismatch in Reynolds numbers. Water tunnels are generally better suited to carrying out fundamental research than they are for applied aerodynamic testing. However, they are very useful as part of a large research program by helping establish the testing schedule for large wind tunnels. In flow situations that are insensitive to Reynolds number, or where a test Reynolds number is close to that of a full-size vehicle, water tunnels should be regarded as the preferred option for experimental aerodynamics. Such examples include micro air vehicles, high-rate dynamic testing, and high-sweep sharp-edge configurations. Water tunnels are also very useful for providing validation data for computational-fluid-dynamics analyses of a flow. An earlier version of this work was prepared for the TTCP TR-AER-TP5 Panel in August 2010.

This research examines the application of using a scanning Light Detection and Ranging(LiDAR) to perform Automated Aerial Refueling(AAR). Speci cally this thesis presents two algorithms to determine the relative position between the tanker and receiver aircraft. These two algorithms require a model of the tanker aircraft and the relative attitude between the aircraft. The rst algorithm ts the measurements to the model of the aircraft using a modi ed Iterative Closest Point (ICP) algorithm. The second algorithm uses the model to predict LiDAR scans and compare them to actual measurements while perturbing the estimated location of the tanker. Each algorithm was tested with simulated LiDAR data before real data became available from test ights. The data collected from this test ight was used to determine the accuracy of the two algorithms with real LiDAR data. After correcting for modeling errors the accuracy of each algorithm is about a Mean Radial Spherical Error of 40cm., The original document contains color images.

The DSTO water tunnel's balance and rotary support mechanism provides a measurement capability for longitudinal dynamic derivatives. This report documents the underlying theory and computational implementation of a technique which uses the water tunnel for determination of normal force and pitching moment coe cient derivatives with respect to angle of attack, non-dimensional pitch rate and angle of attack rate.

Dynamic response analysis tools used by the aerospace industry rely heavily on linear modal frequency response finite element methods. These linear methods are straight forward to use even in the analysis of a complex structural component that require a large number of degrees-of-freedom to model. However, this approach is not suitable for predicting the response of highly loaded thermal/acoustic aircraft structures that may respond in a nonlinear geometric manner. This type of problem requires a nonlinear transient analysis. The nonlinear analysis of a complex structural component is computationally prohibitive, especially for random acoustic response prediction which requires long duration time simulations. To overcome these computational deficiencies, nonlinear reduced order modeling (NLROM) methods have been developed. The objective of this effort is to further refine the NLROM methodology and validate it in the design and lab testing of a representative aircraft structural component. The final product of this program is a validated method and the quantified benefits of using the NLROM methodology., Report contains color.

This report describes the formation of the Navy's Aerodynamics Laboratory and its pioneering contributions in naval aeronautics from 1911 to 1919. The Navy was just beginning serious interest in aviation in 1911 with the procurement of its first aircraft, a Curtiss A-1. This aircraft was technologically similar to the Wright brothers' first airplane but with greater power to allow takeoff from the water using its large central float. At this time the practice of aeronautical engineering was largely a process of trial and error. While this method was successful for small aircraft like the A-l, it posed a significant impediment for development of larger more capable aircraft. Under the leadership of Rear Admiral David W. Taylor, the Navy's Experimental Wind Tunnel was designed and built at the Washington Navy Yard next to the Navy's Experimental Model Basin to advance the state of aeronautical engineering. The Navy's new wind tunnel was the world's largest and the centerpiece of the Navy's Aerodynamics Laboratory. The laboratory, and the naval constructors who worked there under Taylor, developed and refined methods for testing scaled models of complete aircraft as well as aircraft components. These experiments provided the data needed to effectively design large aircraft and led to the success of the Navy's NC flying boat. In 1919, an NC was the first airplane to fly across the Atlantic. This was an accomplishment that at the time was as amazing to the average person as landing on the moon would be fifty years later. In a decade when U.S. advancements in aeronautics were waning, the pioneering work of the Navy's Aerodynamics Laboratory propelled the Navy to the forefront of aeronautics in the second decade of the 20 century. The laboratory established a foundation for the continued development of aeronautics and left a legacy that continues 100 years later., This work was conducted as part of the 2011 celebration of the Centennial Anniversary of Naval Aviation under work unit number 9-2-5300-010-12. It was presented as AIAA paper 2011-6943 at the Centennial of Naval Aviation Forum in Virginia Beach. Virginia, 21-22 September 2011.

This article summarizes the capabilities and development of the Helios version 2.0, or Shasta, software for rotary wing simulations. Specific capabilities enabled by Shasta include off-body adaptive mesh refinement and the ability to handle multiple interacting rotorcraft components such as the fuselage, rotors, flaps and stores. In addition, a new run-mode to handle maneuvering flight has been added. Fundamental changes of the Helios interfaces have been introduced to streamline the integration of these capabilities. Various modifications have also been carried out in the underlying modules for near-body solution, off-body solution, domain connectivity, rotor fluid structure interface and comprehensive analysis to accommodate these interfaces and to enhance operational robustness and efficiency. Results are presented to demonstrate the mesh adaptation features of the software for the NACA0015 wing, TRAM rotor in hover and the UH-60A in forward flight., Presented at AIAA Aerospace Sciences Meeting (49th) including the New Horizons Forum and Aerospace Exposition, Orlando, Florida, Jan. 4-7, 2011 and published as AIAA-2011-1105. The original document contains color images.

Modeling and simulation (M&S) plays an important role when complex human-system notions are being proposed, developed and tested within the system design process. National Aeronautics and Space Administration (NASA) as an agency uses many different types of M&S approaches for predicting human-system interactions, especially when it is early in the development phase of a conceptual design. NASA Ames Research Center possesses a number of M&S capabilities ranging from airflow, flight path models, aircraft models, scheduling models, human performance models (HPMs), and bioinformatics models, among a host of other kinds of M&S capabilities that are used for predicting whether the proposed designs will benefit the specific mission criteria. The Man-Machine Integration Design and Analysis System (MIDAS) is a NASA ARC HPM software tool that integrates many models of human behavior with environment models, equipment models, and procedural / task models. The challenge to model comprehensibility is heightened as the number of models that are integrated and the requisite fidelity of the procedural sets are increased. Model transparency is needed for some of the more complex HPMs to maintain comprehensibility of the integrated model performance. This will be exemplified in a recent MIDAS v5 application model and plans for future model refinements will be presented., See also ADA538937. Presented at the Proceedings of the Conference on Behavior Representation in Modeling and Simulation (19th), held in Charleston, South Carolina, 21 - 24 March 2010, p14-17. Sponsored in part by AFRL, ARI, ARL, DARPA, & ONR.

In this paper we present an application of fault identification and control reconfiguration in the context of a high performance aircraft. A second order divided difference filter is used to identify an internal leakage fault in an electrohydraulic actuator found in the aircraft elevator. The identified fault information is then utilized in the formulation of an aircraft systems model for prognosis-based control. An optimization based reconfiguration strategy is presented to minimize degradation of the fault in presence of performance, actuation, and mission constraints. The strategy is then validated through Hardware-in-the-Loop Simulations., See also ADA539605. European Workshop on Structural Health Monitoring (5th) (EWSHM 2010) held in Naples, Italy on 29 June-J2 uly 2010. U.S. Government or Federal Purpose Rights License. The original document contains color images.

Simulation provides a method of modeling complex systems which would otherwise be impractical for quantitative experimentation. While other analytic techniques have been used to explore Total Non-Mission Capable [due to] Supply (TNMCS) rates, simulation offers a novel approach to discovering what aspects of the supply chain impact this metric. This research develops a discrete event simulation to investigate factors which affect TNMCS rates for the B-1B by modeling the core processes within the Air Force (AF) supply chain. A notional fleet of 16 aircraft at a single air base (Ellsworth AFB, SD) is modeled based on historical supply and maintenance data. To identify and quantify the effects of various factors, an experimental design is used for analyzing the output of our high-level discrete event simulation. Additionally, two different approaches to reporting and modeling Air Logistics Center (ALC) stockage effectiveness (SE) are compared to our baseline simulation. This exploration shows several factors which significantly impact TNMCS rates and have the potential to reduce them to their current targets., The original document contains color images.

There are significantly different manufacturing processes and part counts associated with composites that are not currently addressed within the procurement and life cycle management processes in the Department of Defense (DoD). A series of initiatives have culminated in significant evidence over the last decade to combat the implied risk of composite aircraft. An Air Force Research Labs program, Advanced Composite Cargo Aircraft (ACCA), provides substantial support for the impact of part size on life cycle cost on payload aircraft. This paper evaluates select current methods used and seeks to introduce modifications to the projected manufacturing hours cost. The findings suggest the impact of monolithic parts appears to permeate most of the major cost categories in development and production. The discussion addresses the far-reaching implications of trading several parts for one. This research finds that a significant relationship between relative part count and major cost categories does exist. Specifically, a percentage reduction in part count drives a corresponding percentage reduction in the manufacturing hours. The series of findings pertaining to part count and cost merit consideration for updates to the CERs and interim modifications to capture some portion of the impact in current life cycle cost models., The original document contains color images.

The dynamic derivatives are widely used in linear aerodynamic models which are considered to determine the flying qualities of an aircraft: The ability to predict them reliably, quickly and sufficiently early in the design process is more and more important. This paper describes some experimental and computational activities dealing with the determination of dynamic derivatives. The work has been carried out within the FP6 European project SimSAC. Numerical and experimental results are compared for two aircraft configurations: The generic civil transport aircraft, wing-fuselage-tail configuration DLR-F12 and a generic Transonic CRuiser (TCR), which is a canard configuration. Static and dynamic wind tunnel tests have been carried out for both configurations and are briefly described. The data base generated for the TCR configuration includes small amplitude oscillations, dedicated to the determination of dynamic derivatives, and large amplitude oscillations, in order to investigate the dynamic effects on nonlinear aerodynamic characteristics. The influence of the canard has been investigated. Dynamic derivatives have been determined on both configurations with a large panel of tools, from linear aerodynamic (Vortex Lattice Methods) to CFD (unsteady Reynolds-Averaged Navier-Stokes solvers). Strong limitations of linear aerodynamic tools are observed for the canard configuration. A specific attention is paid to acceleration techniques in CFD methods, which allow the computational time to be dramatically reduced while keeping a satisfactory accuracy., QC 20140815

This report gives details of a five-component strain-gauge balance that has been developed to measure flow-induced loads on models in the DSTO water tunnel. The loads are very small and the balance was designed to measure side and normal forces, as well as rolling, pitching and yawing moments, within the ranges plus or minus 25 N, plus or minus 25 N, plus or minus 0.1 N.m, plus or minus 0.2 N.m and plus or minus 0.2 N.m respectively. These loads are at least 2 to 3 orders of magnitude smaller than those typically measured on aircraft models in the low-speed wind tunnel at DSTO. The balance has not been designed to measure axial forces. Due to the small loads, it was necessary to use semi-conductor strain gauges on the balance. The five-component balance has been developed using similar design principles to a two-component balance developed earlier at DSTO for use in the water tunnel., The original document contains color images.

This report presents a new helicopter transmission model in support of condition-based maintenance (CBM) initiatives. The model uses the first principles approach to analyze the rotating shafts and gear couplings using rotor dynamic analysis techniques. The helicopter transmission is presented, followed by a discussion of the finite element formulation to include gear mesh modeling methodologies. System natural frequencies and mode shapes are discussed, along with the model's validation procedure. The model's primary purpose is to simulate multiple operating conditions to include faults. One such fault condition simulates the wear of gear tooth contact surfaces. The nature of the fault and its modeling techniques are discussed. Simulation results are presented in the time domain using a conventional vibration diagnostic parameter. A new diagnostic parameter, the FE2, is also introduced for detecting faults in the frequency domain. Additional considerations and further research points are also discussed. The results demonstrate the potential of physics-based, mathematical models in providing a key technology enabler for improving health monitoring and prediction capabilities of CBM strategies., The original document contains color images. Prepared in collaboration with Virginia Univ., Charlottesville, VA.

Physics based simulation in conceptual design is widely seen as a way of increasing the information about designs, thus helping with the avoidance of unanticipated problems as the design is refined. This paper reports on an effort to assess the use of CFD level aerodynamics for the development of tables for flight dynamics analysis at the conceptual stage. A number of aerodynamic data sources are used with sampling and data fusion to allow the efficient generation of the tables. A refined design passenger jet wind tunnel model is used as a test case, and three simplified conceptual versions of this geometry are generated. The influence of geometry approximations and modelling influences are evaluated to assess the usefulness of CFD for this application. Finally, the aerodynamic differences are assessed in terms of basic longitudinal flight dynamics analysis., QC 20141001

The objective of this project was to demonstrate and certify a fixed-wing platform for low-cost, high-resolution wide area assessment (WAA) of former and active military facilities contaminated with unexploded ordnance (UXO). The Minimum Altitude Remote Sensing (MARS) airborne UXO mapping system was developed, successfully tested, and deployed in Europe by SeaTerra GmbH in Germany. The principal objectives of this demonstration were to test and evaluate the MARS system in the United States and compare the performance, results, and cost to HeliMag technology. The site selected for demonstration was the Former Kirtland Precision Bombing Range (KPBR), located near Albuquerque, New Mexico. Sky Research, Inc. (SKY) acquired the CT Short Wing (SW) model light sport aircraft (LSA) for this demonstration. This model lightweight aircraft, a slightly different model than the one used by SeaTerra in Europe, did not affect the deployment of the technology. The aircraft has an 8.9 meter wingspan, a length of 6.2 meters, and can fly at a minimum height above ground of 2 meters. It uses modern German glider structural design techniques and is essentially all-composite glass fiber reinforced plastic with very little ferrous metal and a low signature footprint. An array of six Geometrics Model 822A cesium vapor magnetometer sensors was mounted on a sensor boom attached to the aircraft wings. The sensors were deployed at separations between 1.15 and 2.0 meters. SeaTerra's AGS MK3 frequency counter and data acquisition system (DAS) were used to record data at between 140 to 160 Hertz (Hz) with a ~0.2 nanotesla (nT) noise floor on the outer sensors and 0.5-0.7 nT noise floor on the inner sensors (after processing). Daily flight durations were planned to be approximately 6 hours per day (assuming an 8-hour daily window for flying)., The original document contains color images.

Due to the capability shortfalls associated with the transformation of the Marine Corps aviation force structure and the challenges associated with a changing security environment, the role of the utility helicopter must evolve to achieve a true multi-mission capability. Through a change in paradigm, the Marine Corps must properly equip the UH-lY to meet these challenges. Conclusion: The current composition of the Marine Corps aviation force structure will remain constant for the next 25 years. The transition of upgraded type model series aircraft within the existing force structure coupled with the constraints of a bureaucratic acquisitions process and the security challenges associated with irregular warfare, security cooperation initiatives and the long war have generated significant capabilities gaps within its operational requirements. The solution will therefore require a timely and fiscally responsible alternative that will involve a paradigm change within existing helicopter doctrine and employment. That solution is the evolution of the utility helicopter into a true multi-role aircraft that possesses the capability, equipment, and versatility to meet the challenges of a changing security environment.

To characterize and improve upon aircraft designs in an expedient and efficient manner, it is imperative to generate approaches for rapidly assessing new aircraft designs. Recent technological advancements have enabled faster and more accurate assessment in three specific areas: computational fluid dynamics (CFD) modeling, rapid prototyping (RP), and experimental global measurements. The Air Force Research Laboratory Air Vehicles Directorate (AFRL/RB) initiated a joint computational/experimental ground testing program to investigate and analyze the flow field of the unmanned combat air vehicle (UCAV) X45-A as well as a strike tanker configuration. The test program used various rapid prototyping manufacturing technologies to fabricate models for ground testing; implemented innovative, nonintrusive measurement techniques; and compared CFD results with experimental data acquired from ground tests., Prepared in cooperation with the Advanced Sensing & Optical Measurement Branch, National Aeronautics and Space Administration, Hampton, VA and Innovative Scientific Solutions, Inc., Dayton, OH.

In view of a continuous increase of computer performance, it is nowadays feasable to use CFD (Computational Fluid Dynamics) analysis very early in the conceptual design stage, if not the pre-concept phase, of aircraft. This requires the generation of a CAD (Computer Aided Design) model suitable for CFD computations, which is a tedious and time consuming process because a disconnect exists between the geometrical definition required for a CAD model and the limited number of geometry related design parameters defined by the designer. An additional complication to this is the requirement of producing a closed and consistent CAD model suitable for problem setup of CFD computations. The CADac (CAD-aircraft) tool nils this gap by automating the generation of closed and consistent CAD models via the implementation of a parameterized approach to conceptual design. CADac enables therefore to use CFD earlier and to use tools with inter-laced fidelity at the conceptual design phase., QC 20101104

Distributed Mission Operations (DMO) is an ideal setting for practicing Beyond Visual Range air-to-air tactics. Hardware and software limitations often dictate the use of simplified aerodynamic models for control of fixed wing constructive entities within synthetic environments. In many tactical situations the long range fight will disintegrate into close-in air combat, which for a variety of reason is difficult to represent in virtual simulators. The eXperimental Common Immersive Theatre Environment (XCITE) developed at Warfighter Readiness Research Division (AFRL/HEA) was designed to provide a physics based high-fidelity threat environment for training and rehearsal in a DMO environment. Feedback by operational pilots identified unrealistic constructive flight performance as a critical shortfall of XCITE. Here we will discuss the development, testing and validation of an energy based aerodynamic model in an effort to provide a more effective threat environment for DMO. Specific Excess Power (Ps) tables were created using thrust, drag, lift and loading data for each aircraft; these tables then provide accurate acceleration figures that are feed back into the original aero-model. In testing, aircraft flight performance was compared against data obtained from aerodynamic models in both 6-DOF Full Mission Trainers and the baseline XCITE aero model., Presented at Spring Simulation Multiconference, SpringSim 2007, Norfolk VA on 29-29 Mar 07, and published in proceedings of the same, v3 p327-336. ISBN 1-56555-314-4. The original document contains color images.

This paper presents a novel topology optimization methodology for a synthesis of distributed actuation systems with specific applications to morphing air vehicle structures. The main emphasis is placed on the topology optimization problem formulation and the development of computational modeling concepts. The analysis model is developed to meet several important criteria: It must allow a large rigid-body displacement, as well as a variation in planform area, with minimum strain on structural members while retaining acceptable numerical stability for finite element analysis. For demonstration purposes, the in-plane morphing wing model is presented. Topology optimization is performed on a semiground structure with design variables that control the system configuration. In other words, the state of each element in the model is controlled by a corresponding design variable that, in turn, is determined through the optimization process. In effect, the optimization process assigns morphing members as soft elements, nonmorphing load-bearing members as stiff elements, and nonexistent members as "voids." The optimization process also determines the optimum actuator placement, where each actuator is represented computationally by equal and opposite nodal forces with soft axial stiffness. Sample in-plane morphing problems are solved to demonstrate the potential capability of the methodology., Published in Journal of Aircraft, v44, n4, p1205-1213, Jul-Aug 2007.

The Air Force Research Laboratory Munitions Directorate (AFRL/MN) is presently involved in many aspects of micro aerial vehicle (MAV) research. Among these are: advanced modeling and simulation models for MAVs, aero-structural interaction, advanced guidance techniques, hardware-in-the-loop simulations, and vehicle integration. In order to optimize collaboration within AFRL and also with outside research organizations, it was decided that a common MAV configuration be designed that would serve as a reference for current and future research. This paper describes a generic micro air vehicle that will serve as a "baseline" configuration. The MAV design incorporates a circular fuselage, a thin cambered wing, and a conventional tail. The MAV has a wingspan of 24 inches and a fuselage length of 17 inches. This paper will also detail the rational behind the design as well as provide initial aerodynamic properties and flight performance characteristics of the AFRL Generic MAV, herein called "GENMAV.", Prepared in collaboration with the Department of Mechanical Engineering, Oregon State University, Corvallis, OR. Presented at the AIAA Aerospace Sciences Meeting (45th) held in Reno, NV, 8-11 Jan 2007. Published as AIAA 2007-0667. The original document contains color images.