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In the agricultural aviation, there are several aerodynamic factors that must be optimized in order to contribute to the successful application of agricultural products, such as the high aerodynamic efficiency (L/D) required at the working phase and the influence of aircraft wingtip phenomena on the spray deposition and movement. For these reasons, in this research is presented the conceptual design of an advanced prototype of agricultural aircraft, whose main characteristic is an adaptive multi-winglet device installed on the wingtips, which optimized the main aerodynamic issues presented in its mission. Traditional aircraft design methods were used to develop and assess the suitability of the aircraft, focusing on its design requirements and tackling studies of weight sizing, pilot ergonomics, aerodynamics, stability, and performance. Subsequently, analytical and computational methods were used to design the adaptive multi-winglet device, which is composed by three winglets with its own geometry fitted on a tip-tank. Six configurations were created by modifying only the cant angle of each winglet in order to determine the arrangement that provides the best aerodynamic characteristics through a study of computational fluid dynamics (CFD), using the Reynolds–Averaged–Navier–Stokes (RANS) equations coupled with the Shear Stress Transport (SST) turbulence model. First, the flow around the wing and the multi-winglet section of the aircraft was investigated exclusively. Afterward, the airflow around the entire aircraft was studied at the product application condition, in order to compare the overall aerodynamic performance of the baseline concept along with the optimal multi-winglet configuration installed on the aircraft. Lift, drag and pitching moment coefficients were assessed, as well as the wingtip vortex structure of the most relevant configurations. Results of this study showed that adaptive multi-winglet devices are a promising alternative to improve the overall, QC 20181023

Computational Fluid Dynamics (CFD) has become a major tool in aerodynamic analysis throughout the aerospace industries, complementary to traditional methods such as wind tunnel testing, and analytical calculations. In this research, an attempt was made to integrate the Similarity and Area Rules with CFD methods. Both tools, the Similarity/Area Rule and CFD are used to derive the characteristics of complicated aerodynamic shapes in the transonic Mach number regime. It was found that the Similarity Rule can only be verified qualitatively. On the other hand, the Area Rule can be more completely verified. The aim was to find ways to minimize the drag of the tralifrig configurations of the Arr-to-Ground (A/G) weapon, Joint-Standoff-Weapon GSO%Q), in its Captive- Air-Training-Missile (CAm4) configuration. By analyzing the combination of CAmI and Pylon, it was found that the drag of this configuration depends on the average slope of the area cross-section distribution of the afterbody. The CFD tools used were a state-of-the-art grid generation code, GRIDGEN, and a multi- grid integration code, PEGSUS; the configurations were run with the OVERFLOW solver using Euler, as well as Navier-Stokes solutions. For drag optimization, Euler solutions give adequate results, the need for NS solution can be restricted to more intensity viscous analysis, http://archive.org/details/aerodynamicnalys109457859, Major, Israeli Air Force, Approved for public release; distribution is unlimited.

This report is a User's Manual for the 2011 Revision of the Missile Datcom computer program. It supersedes AFRL-RB-WP-TR-2009-3015. In missile preliminary design it is necessary to quickly and economically estimate the aerodynamics of a wide variety of missile configuration designs. Since the ultimate shape and aerodynamic performance are so dependent upon the subsystems utilized, such as payload size, propulsion system selection and launch mechanism, the designer must be capable of predicting a wide variety of configurations accurately. The fundamental purpose of Missile Datcom is to provide an aerodynamic design tool which has the predictive accuracy suitable for preliminary design, and the capability for the user to easily substitute methods to fit specific applications., Supercedes 2008 Revision, AFRL-RB-WP-TR-2009-3015, ADA503576.

The High Enthalpy Shock Tunnel Goettingen (HEG) of the German Aerospace Center (DLR) is one of the major European hypersonic test facilities. It was commissioned for use in 1991 and was utilised since then extensively in a large number of national and international space and hypersonic flight projects. Originally, the facility was designed for the investigation of the influence of high temperature effects such as chemical and thermal relaxation on the aerothermodynamics of entry or re-entry space vehicles. Over the last years its range of operating conditions was subsequently extended. In this framework the main emphasis was to generate test section conditions which allow investigating the flow past hypersonic flight configuration from low altitude Mach 6 up to Mach 10 in approximately 33 km altitude. The studies performed focused on the external as well as internal aerodynamics including combustion of hydrogen in complete supersonic combustion ramjet configurations. The complexity of these flows requires that experiments in ground based facilities are strongly linked with computational fluid dynamics (CFD) investigations. These common activities range from the calibration process of the facility and the study of basic aerodynamic configurations, which are well suited to look at fundamental aspects of high enthalpy flow fields, to the investigation of complex re-entry configurations. In the DLR Spacecraft Department the research programs conducted in HEG are closely linked with numerical studies using the DLR TAU code., See also ADA569031. Aerothermodynamic Design, Review on Ground Testing and CFD (Conception aerothermodynamique, revue sur les essais au sol et dynamique des fluides informatisee). Presented at the AVT-186 RTO AVT/VKI Lecture Series, von Karman Institute, Rhode St. Genese, Belgium, 29 March-1 April 2010, paper 8.

The flow past and loads on an unmanned combat air vehicle (UCAV) 1303 was investigated in a water tunnel. The UCAV 1303 is characterized by a non-slender, 47-deg swept-leading-edge delta wing, with a cranked trailing edge, a cropped wing tip and a fuselage. It has no vertical tail. Dye visualization and a five-axis strain gage load data were obtained for both steady and unsteady flows, from which body axis force and moment data were derived. Comparison of the flow with slender delta wings showed a number of remarkable differences such as the presence of two like-sense vortices on the same side of the wing, their simultaneous spiral and bubble bursting; as well as some similarities with the slender delta wing behavior. This paper discusses these and also compares the steady flow and unsteady flow results., See also ADA540154. Royal Aeronautical Society Aerodynamics Conference 2010. Applied Aerodynamics: Capabilities and Future Requirements. Held in Bristol, United Kingdom on July 27-28, 2010. U.S. Government or Federal Purpose Rights License. The original document contains color images.

A Lagrangian viscous vortex technique is used in the present study to simulate the unsteady flow field of flapping flight. The method is grid free and computations are done only at the regions of non-zero vorticity. As a result this method is quite suitable for simulating the unsteady vertical flow-field past flapping wings and airfoils; it provides an accurate and fast flow-visualization. The main strength of this tool is its grid free nature; as the resolution of the grid is a crucial parameter in resolving the unsteady flow-field accurately. It is also relatively easy to implement a grid-free particle based technique for a moving boundaries or complex geometries. In this algorithm, field property vorticity is carried by discrete particles which convects with the Biot-Savart velocity and diffuses with random walk.

To address the complex multidisciplinary nature of rotorcraft analysis, high-fidelity computational fluid and structural dynamics models have been developed to investigate a range of challenging rotorcraft issues. First, an advanced technology, active flap rotor (Boeing SMART) is investigated, and performance, aerodynamic and structural loads, vibration, noise prediction and flow physics mechanisms are shown. The rotor model includes complex and detailed flap and flap gap modeling. Second, analyses on an advanced dynamics model (ADM) research configuration rotor investigate regressing lag mode (RLM) aeroelastic instabilities. Tightly-coupled CFD/CSD stability calculations show noticeable improvement over lower fidelity methods. Third, the state-of-the-art capability of CFD methods to directly predict low frequency in-plane noise on realistic lifting rotors is benchmarked for the first time. In all cases, comparisons are made between CFD/CSD, comprehensive analyses, and experimental data. Taken together, these works offer an important advancement in rotorcraft analysis capability for advanced technology rotor configurations under study for future Army rotorcraft and highlight future needs in next generation rotorcraft analysis software., Presented at the DoD High Performance Computing Modernization Program (HPCMP) (20th) Users Group Conference (UGC), 2010. Prepared in cooperation with Eloret Corp., Moffett Field, CA. The original document contains color images.

The new goal for unmanned aerial systems will be to find creative methods of keeping the cost low and still maintain effectiveness. This thesis discusses the importance of UAVs over the last few years, suggests the development of a low-cost, large UAV, and evaluates the results. We also examine the idea of a platform for deploying multiple aerial-delivery, parafoil-based systems and discuss scenarios for the improvement of the collaboration of the large UAV with the Snowflake project., The original document contains color images.

A slowed-rotor compound helicopter has been synthesized using the NASA Design and Analysis of Rotorcraft (NDARC) conceptual design software. An overview of the design process and the capabilities of NDARC are presented. The benefits of trading rotor speed, wing-rotor lift share, and trim strategies are presented for an example set of sizing conditions and missions., Presented at the American Helicopter Society Aeromechanics Specialist's Conference, San Francisco, CA, 20-22 Jan 2010. Prepared in collaboration with the National Aeronautics and Space Administration (NASA) Ames Research Center, Moffett Field, CA.

Coupling of aeromechanics analysis with vehicle sizing is demonstrated with the CAMRAD II aeromechanics code and NDARC sizing code. The example is optimization of cruise tip speed with rotor/wing interference for the Large Civil Tiltrotor (LCTR2) concept design. Free-wake models were used for both rotors and the wing. This report is part of a NASA effort to develop an integrated analytical capability combining rotorcraft aeromechanics, structures, propulsion, mission analysis, and vehicle sizing. The present paper extends previous efforts by including rotor/wing interference explicitly in the rotor performance optimization and implicitly in the sizing., Presented at the American Helicopter Society Aeromechanics Specialist's Conference, San Francisco, CA, 20-22 Jan 2010.

This paper overviews the SimSAC Project, Simulating Aircraft Stability And Control Characteristics for Use in Conceptual Design. It reports on the three major tasks: Development of design software, validating the software on benchmark tests and applying the software to design exercises. CEASIOM, the Computerized Environment for Aircraft Synthesis and Integrated Optimization Methods, is a framework tool that integrates discipline-specific tools for conceptual design. At this early stage of the design it is very useful to be able to predict the flying and handling qualities of this design. In order to do this, the aerodynamic database needs to be computed for the configuration being studied which then has to be coupled to the stability and control tools to carry out the analysis. The benchmarks for validation are the F12 windtunnel model of a generic long-range airliner and the TCR windtunnel model of a sonic-cruise passenger transport concept. The design, simulate and evaluate (DSE) exercise demonstrates how the software works as a design tool. The exercise begins with a design specification and uses conventional design methods to prescribe a baseline configuration. Then CEASIOM improves upon this baseline by analyzing its flying and handling qualities. Six such exercises are presented., QC 20201117

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

The report contains a study of the phenomenon of leading-edge vortex breakdown over slender delta wings conducted by the NATO/RTO/AVT Task Group AVT-080. After an introduction and discussion of highlights, a chapter is devoted to a critical assessment of available experimental studies of the vortex breakdown phenomenon. Eight reliable experimental data sets are collected and compiled. These data sets are described in separate chapters and were obtained on more or less identical delta wing models tested in different wind tunnel test facilities. The datasets include timeaverage results at static model conditions, ensemble-average results at dynamic model conditions and unsteady results at static/dynamic model conditions. Based on the critical assessment, a benchmark test case is selected that provides sufficiently reliable data for validation of the numerical predictions. In separate chapters ten state-of-the-art numerical solutions from various research institutes are described that have been obtained for the benchmark test case configuration and test conditions. Studies of the effects of parameters like grid density, turbulence modeling, time scaling, etc., have been performed. In addition, for a special delta wing configuration, the effects of tunnel wall interference and of interactions between vortices have been calculated. In the final three chapters advanced analytical methods are described. The report is of great value to scientists involved in the development of design methods for military aircraft configurations, but also for high speed civil and/or space aircraft., See also ADM002319., The original document contains color images. ISBN978-92-837-0078-4

A cooperative study was performed to investigate the handling quality effects from reduced flight control system stability margins, and the trade-offs with higher disturbance rejection bandwidth (DRB). The piloted simulation study, performed on the NASA-Ames Vertical Motion Simulator, included three classes of rotorcraft in four configurations: a utility-class helicopter; a medium-lift helicopter evaluated with and without an external slung load; and a large (heavy-lift) civil tiltrotor aircraft. This large aircraft also allowed an initial assessment of ADS-33 handling quality requirements for an aircraft of this size. Ten experimental test pilots representing the U.S. Army, Marine Corps, NASA, rotorcraft industry, and the German Army (WTD 61), evaluated the four aircraft configurations, for a range of flight control stability margins and turbulence levels, while primarily performing the ADS-33 Hover and Lateral Reposition MTEs. Pilot comments and aircraft-task performance data were analyzed. The preliminary stability margin results suggest higher DRB and less phase margin cases are preferred as the aircraft increases in size. Extra care will need to be taken to assess the influence of variability when nominal flight control gains start with reduced margins. Phase margins as low as 20-23 degrees resulted in low disturbance response damping ratios, objectionable oscillations, PIO tendencies, and a perception of an incipient handling qualities cliff. Pilot comments on the disturbance response of the aircraft correlated well to the DRB guidelines provided in the ADS-33 Test Guide. The ADS-33 mid-term response-to-control damping ratio metrics can be measured and applied to the disturbance-response damping ratio. An initial assessment of LCTR yaw bandwidth shows the current Level 1 boundary needs to be relaxed to help account for a large pilot off-set from the c.g., Presented at the American Helicopter Society Annual Forum (65th), Grapevine, Tx, 27-29 May 2009.

This report results from a contract tasking Institute of Theoretical and Applied Mechanics as follows: At sufficiently high angles of attack the boundary layer globally separates from the leeward surface of delta wing. The flow field contains two strong vortices generated by the roll up of the shear layer emanating from the separation lines near the wing leading edges. For supercritical angles, the vortices demonstrate a large scale instability accompanied by strong oscillations and eventual breakdown of the vortical structure. This phenomenon, called vortex breakdown or vortex burst, leads to rapid decrease of the lift to drag ratio. If the vortex burst on one wing side occurs earlier than on the other, then strong yawing and rolling moments are induced due to flow asymmetry. It is well known that the leading-edge vortex flow is extremely sensitive to variations of the surface shape and/or external forcing produced near the separation line. Such forcing can be organized by means of dielectric barrier discharge actuators which may be advantageous compared to other modern control techniques (synthetic jets, MEMS and etc.). Dielectric barrier discharge has simple construction, does not change the aerodynamic shape or influence the wing functionality when it is not in use, allows for smooth variations of forcing frequency and power, and can be used for closed loop feedback control. This kind of activation was successfully used for control of boundary-layer separation near the leading edge of a subsonic airfoil at sufficiently high angles of attack. The problem will be studied experimentally in the ITAM low-speed wind tunnel. The experiments will be performed on a delta wing model of approximately 0.5 m chord for flow speeds 10 30 m/s. The model will simulate flow features of the generic ?Stingray? UAV model tested with the synthetic-jet actuators. Tests will be conducted at chord Reynolds numbers up to 1 million , which are typical for UAV applications., The original document contains color images.

This project investigated several areas with the aim of improving performance and operational use of nano air vehicles (NAVs) by understanding fundamental operating principles and developing key technologies. The project developed plasma actuators for thrust and maneuvering control and examined flapping flight through the development of a high speed videogrammetry system and a gust/shear tunnel to replicate flow fields seen as the NAV scale. In the first portion of the research effort, we developed plasma actuators for use as control surfaces, stall mitigation devices and propulsive units for NAVs. Both of these concepts leverage the relevant low speed flow physics experienced at the NAV scale. A secondary objective of this project was to investigate the physics of low speed flow over various biologically inspired aerodynamic configurations, from both numerical and experimental perspectives, and use this information to develop novel yet practical concepts for NAVs. The primary focal points of biological inspiration for this project were the exploration of low speed aerodynamics - particularly separation phenomena - and the physical mechanisms used by nature to optimize performance in this flight regime, including the use of novel plasma actuators, aeroelastic behavior and its effect on gust response and alleviation, and the impact of unsteady motions (flapping) on the flow, especially for the reduction of drag and separation. This was related to the primary objectives of the proposal since plasma actuators are particularly useful as laminar separation control devices., Sponsored in part by DARPA. The original document contains color images. All DTIC reproductions will be in black and white.

The trade-off between accuracy and speed is a re-occurring dilemma in many facets of military performance evaluation. This is an especially important issue in the world of ISR. One of the most progressive areas of ISR capabilities has been the utilization of Unmanned Aircraft Systems (UAS). Many people believe that the future of UAS lies in smaller vehicles flying in swarms. We use the agent-based System Effectiveness and Analysis Simulation (SEAS) to create a simulation environment where different configurations of UAS vehicles can process targets and provide output that allows us to gain insight into the benefits and drawbacks of each configuration. Our evaluation on the performance of the different configurations is based on probability of correct identification, average time to identify a target after it has deployed in the area of interest, and average time to identify all targets in an area., The original document contains color images.

A computational research program was performed at Georgia Institute of Technology, Penn State University, and at Northern Arizona University to develop a set of first-principles based computational modeling tools for analyzing and designing advanced helicopter configurations. The approach involved incorporation of advanced numerical algorithms and turbulence models in OVERFLOW 2, development of advanced comprehensive analyses (DYMORE and RCAS) that are seamlessly coupled to the flow analysis, modeling of rotor noise characteristics using an advanced acoustics prediction tool (PSU-WOPWOP) that is seamlessly coupled to the flow analysis and the comprehensive analyses, and validation and application of the integrated suite of tools for current generation (UH-60, BO105) and next generation configurations. Under the Phase I-B extension, assessment of this suite of tools is being performed by Ga Tech and Penn State for the Boeing MD-900 model rotor (MDART), an actively controlled rotor (SMART), and the Comanche rotor blade (as an option)., Prepared in cooperation with Pennsylvania State University, University Park, PA. The original document contains color images.

This report is a User's Manual for the 2008 Revision of the Missile Datcom computer program. It supersedes AFRL-VA-WP-TR-1998-3009. In missile preliminary design it is necessary to quickly and economically estimate the aerodynamics of a wide variety of missile configuration designs. Since the ultimate shape and aerodynamic performance are so dependent upon the subsystems utilized, such as payload size, propulsion system selection and launch mechanism, the designer must be capable of predicting a wide variety of configurations accurately. The fundamental purpose of Missile Datcom is to provide an aerodynamic design tool which has the predictive accuracy suitable for preliminary design, and the capability for the user to easily substitute methods to fit specific applications., This document supersedes ADA344707.

Fuel efficiency of aircraft is of great importance to the military and private sector. A more efficient wing design for UAVs would lead to improvements in mission support while reducing fuel costs for the Air Force. An experimental investigation of one candidate design, the Houck Aircraft Configuration, has been conducted in the AFIT low speed wind tunnel. This aircraft shares similarities to other joined-wing aircraft, but includes curved flow-guides of varying spanwise camber connecting the upper and lower wingtips. Experimental results that the addition of flow guides on the 24" Houck Configuration results in a 2.5% reduction in L/D(sub max) at Re =80K and a 0.3% reduction in L/Dmax at Re= 125K. This trend shows a decrease in the performance gap as the Reynolds number increases from 80K to 125K. It is recommended that additional testing at higher Reynolds numbers be performed to determine if an increase in performance can be shown. The designed flow guides proved to be successful in combining the upper and lower wing-tip vortices into a single vortex. The flow guides alter what would be two smaller compact vortices and instead produce a slightly larger, spread out vortex which follows the curve of the flow guide. Ultimately, evidence of improvements in aerodynamic efficiency will need to be shown before other claims of the design are demonstrated to be fully successful., The original document contains color images.

Radio frequency based discharges at atmospheric pressures are the focus of increased interest in aerodynamics because of the wide range of potential applications including, specifically, actuation in flows at moderate speeds. Recent literature describing promising experimental observations, especially on separation control, has spurred efforts in the development of parallel theoretical modeling to lift limitations in the current understanding of the actuation mechanism. The present effort demonstrates higher fidelity first-principle models in a multidimensional finite-element framework to predict surface discharge-induced momentum exchange. The complete problem of a dielectric barrier discharge at high pressure with axially displaced electrodes is simulated in a self-consistent manner. Model predictions for charge densities, the electric field, and gas velocity distributions are shown to mimic trends reported in the experimental literature. Results show that a residual of electrons remains deposited on the dielectric surface downstream of the exposed powered electrode for the entire duration of the cycle and causes a net electric force in the direction from the electrode to the downstream surface. For the first time, results document the mitigation process of a separation bubble formed due to flow past a flat plate inclined at 12 angle of attack. This effort sets the basis for extending the formulation further to include polyphase power input in multidimensional settings, and to apply the simulation method to flows past common aerodynamic configurations., Published in Physics of Plasmas, v13 p023503-1/023503-11, 2006.

An investigation on side jets in an M = 5.0 flow was conducted on a flat plate as well as on a generic missile configuration. Surface pressure measurements, oil flow visualisations, force measurements and associated numerical simulations gave insight into the interaction of different jet configurations and their influence on the efficiency of side jet control described by the amplification factor. The analysis of the data especially in the wake of the jet showed that the low pressure region is much less dominant than assumed and that the position of the side jet nozzle is vital also with respect to the exploitation of a desired amplification of the side jet thrust., See also ADM401233. RTO-MP-AVT-135, Presented at the RTO Applied Vehicle Technology Panel (AVT) Business Meeting Week in Amsterdam, the Netherlands, 15-18 May 2006, AC/323(AVT-135)TP/104. The original document contains color images.

A mission program named Extended Area Protection and Survivability was initiated at the U.S. Army Research Laboratory to demonstrate guided ammunition technologies to defend the battle space against any presented targets, such as mortars, rockets, and artillery. This report introduces the first step in the development of the guided ammunition system. It presents high-level physics-based simulations of a 60-mm projectile system. The topology of the initial projectile was developed on the basis of gun barrel specifications and certain aerodynamics characteristics. A three-dimensional finite element model was created to present the projectile system. Three different projectile configurations were analyzed with the LS-DYNA program and the results were compared. An acceptable deign demonstrated that the muzzle velocity reached only 85% of target value mainly because the launch package exceeded the desired mass by 25%. In addition, a characteristic centerline variation of a gun barrel was taken into account in this report. The outcome of the in-bore vertical displacements and accelerations attributable to the variation was found to be significant. Note that the structural configuration of the initial projectile is not optimal. Rigorous optimization efforts will be made on the system, particularly on the sabot component, in order to meet performance requirements., The original document contains color images.

Radio frequency based discharges at atmospheric pressures are the focus of increased interest in aerodynamics because of a wide range of potential applications including specifically actuation in flows at moderate speeds. Recent literature describing promising experimental observations, especially on separation control, have spurred efforts in the development of parallel theoretical modeling to alleviate limitations in current understanding of the actuation mechanism. The present effort builds on a recently developed finite element-based one- and two-dimensional multi-fluid formulation of plasma sheath for atmospheric optical glow discharge in partially ionized gas. The model was relatively straightforward but formed the foundation of a versatile first-principles based methodology. Higher-fidelity models are included to yield a more sophisticated framework to predict discharge-induced momentum exchange. Here, the complete problem of a dielectric barrier discharge based separation control with axially displaced electrodes is simulated in a self-consistent manner. Model predictions for transient evolution of charge densities, the electric field, electrodynamic force and induced gas velocity distributions for helium gas in quiescent condition are shown to mimic trends reported in the experimental literature. For the first time, results also document the decay process of a separation bubble formed due to flow past a flat plat inclined at 12 deg angle of attack. This effort sets the basis for extending the formulation further to include polyphase power input in multi-dimensional settings, and to apply the simulation method to flows past common aerodynamic configurations., Presented at the AIAA Fluid Dynamics Conference (35th) and AIAA Plasma Dynamics and Lasers Conference (36th) held in Toronto, Canada on 6-9 June 2005. Supported in part by AFOSR.

The Air Force Research laboratory, Space Vehicles Directorate (AFRL/VS), and CSA Engineering have developed a large scale structural testing facility on Kirtland AFB, NM. This facility is capable of applying static and dynamic loads with up to 18 independent hydraulic actuators. Coupled with the servo-hydraulic load control unit is a fully integrated 256-channel data acquisition system (DAS). Configurable for strain gages, LVDT's, or virtually any other strain gage-based or high level sensor. The DAS and load controller communicate and function simultaneously through software written by MTS Corporation. To date, AFRL and CSA Engineering have conducted four major static tests on developing launch vehicle technologies. These tests include two payload adapters, a large advanced grid stiffened composite fairing, and an advanced Payload Attach Fitting (PAF). Additional challenges include accommodating wide ranging physical geometries and applying loads to these geometries. Detailed test procedures, test design, and a testing overview for each static test configuration will be discussed., Pub. in International SAMPE Technical Conference p769-783, 2004.

Viewgraphs only of Membrane Wing Aerodynamics for micro AV Applications., See also ADM001685. Published in CSP 02-5078. Presented at Proceedings for Aerodynamic Issues of Unmanned Air Vehicles (UAV) held in Bath, United Kingdom on 4-5 Nov 2002. The original document contains color images.