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The paper shows how the use of an instrument with 12 small-diameter (47-mm) membrane filters (samples taken for 12 2-hr periods), in conjunction with wind direction measurements, is a suitable technique for observing mass loading levels on a time-resolved basis. In contrast, standard high-volume instruments are unable to demonstrate specialized events because of the long collection period involved. A comparison is presented for a typical day where the mass loading changed during the observed period but could not be observed with a single high-volume sample measurement. Time-averaged mass loading for 2-hr periods is demonstrated to be a useful tool in developing trend analysis for diurnal effects.
Transverse filament tape (TFT) fiberglass/epoxy and TFT polyester composites intended for low cost wind turbine blade fabrication have been subjected to static and cyclic load behavior tests whose results are presently evaluated on the basis of an integrated hygrothermomechanical response theory. Laminate testing employed simulated filament winding procedures. The results obtained show that the predicted hygrothermomechanical environmental effects on TFT composites are in good agreement with measured data for various properties, including fatigue at different R-ratio values.
Three aspects of the test results are compared: rotor blade bending loads, rotor teeter response, and nacelle yaw moments. As a result of the tests, it is shown that while mean flatwise bending moments were unaffected by the placement of the rotor, cyclic flatwise bending tended to increase with wind speed for the downwind rotor while remaining somewhat uniform with wind speed for the upwind rotor, reflecting the effects of increased flow disturbance for a downwind rotor. Rotor teeter response was not significantly affected by the rotor location relative to the tower, but appears to reflect reduced teeter stability near rated wind speed for both configurations. Teeter stability appears to return above wind speed, however. Nacelle yaw moments are higher for the upwind rotor but do not indicate significant design problems for either configuration.
The debate over when wind-tunnel testing will be replaced by Computational Fluid Dynamics (CFD) comes and goes. More recently the debate has subsided with a more collaborative spirit between practitioners of these two disciplines resulting in significant improvements in the outcomes of both. There may come a time, however, when CFD has sufficient accuracy to supplant WTT as the dominant or perhaps only tool for aerodynamic simulation. If and/or when that happens, financial pressures favor efforts to close or severely limit the operations of wind tunnels. Presumably additional resources will go toward CFD to generate aerodynamic databases, load environments, and new aero/fluid-dynamic knowledge. It is therefore important to develop appropriate processes by which wind-tunnel closure decisions are made to ensure that facilities critical to industry and government research and development aren’t closed prematurely without proof that the available CFD tools have sufficient accuracy and low-enough cost (and enough experts and computational facilities) to take on the traditional role of wind tunnels. This paper will describe a project intended to answer the specific question of whether CFD can replace wind-tunnel testing for the limited Mach-number range 2.4 to 4.6. The project involves wind-tunnel testing and coordinated CFD for a variety of vehicle and flow-physics types in the high-speed leg of the Unitary Plan Wind Tunnel at NASA’s Langley Research Center.
A full-scale wind tunnel test was recently conducted (March 2009) in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-FootWind Tunnel to evaluate the potential of an individual blade control (IBC) system to improve rotor performance and reduce vibrations, loads, and noise for a UH-60A rotor system [1]. This test was the culmination of a long-termcollaborative effort between NASA, U.S. Army, Sikorsky Aircraft Corporation, and ZF Luftfahrttechnik GmbH (ZFL) to demonstrate the benefits of IBC for a UH-60Arotor. Figure 1 shows the UH-60Arotor and IBC system mounted on the NFAC Large Rotor Test Apparatus (LRTA). The IBC concept used in the current study utilizes actuators placed in the rotating frame, one per blade. In particular, the pitch link of the rotor blade was replacedwith an actuator, so that the blade root pitch can be changed independently. This concept, designed for a full-scale UH-60A rotor, was previously tested in the NFAC 80- by 120-FootWind Tunnel in September 2001 at speeds up to 85 knots [2]. For the current test, the same UH-60A rotor and IBC system were tested in the 40- by 80-FootWind Tunnel at speeds up to 170 knots. Figure 2 shows the servo-hydraulic IBC actuator installed between the swashplate and the blade pitch horn. Although previous wind tunnel experiments [3, 4] and analytical studies on IBC [5, 6] have shown the promise to improve the rotor s performance, in-depth correlation studies have not been performed. Thus, the current test provides a unique resource that can be used to assess the accuracy and reliability of prediction methods and refine theoretical models, with the ultimate goal of providing the technology for timely and cost-effective design and development of new rotors. In this paper, rotor performance and loads calculations are carried out using the analyses CAMRAD II and coupled OVERFLOW-2/CAMRAD II and the results are compared with these UH-60A/IBC wind tunnel test data.
The results of a wind tunnel test program to determine the performance loads and dynamic characteristics of the Composite Flexbeam Tail Rotor (CFTR) for the AH-64 Advanced Attack Helicopter are reported. The CFTR uses an elastomeric shear attachment of the flexbeam to the hub to provide soft-inplane S-mode and stiff-inplane C-mode configuration. The properties of the elastomer were selected for proper frequency placement and scale damping of the inplane S-mode. Kinematic pitch-lag coupling was introduced to provide the first cyclic inplane C-mode damping at high collective pitch. The CFTR was tested in a wind tunnel over the full slideslip envelop of the AH-64. It is found that the rotor was aeroelastically stable throughout the complete collective pitch range and up to rotor speeds of 1403 rpm. The dynamic characteristics of the rotor were found to be satisfactory at all pitch angles and rotor speeds of the tunnel tests. The design characteristics of the rotor which permit the high performance characteristics are discussed. Several schematic drawings and photographs of the rotor are provided.
Results of tests conducted using the MOD-0 100 kW experimental wind turbine are evaluated. The teetered rotor significantly decreased loads on the yaw drive mechanism and reduced blade cyclic flapwise bending moments by 25 percent at the 20 percent span location when compared to the rigid hub rotor. The teetered hub performed well, but impacted the teeter stops on occasion as wind speed and/or direction varied rapidly. The tip-controlled rotor performed satisfactorily with some expected loss of control when compared to the full span pitchable blade. The performance results indicate that a review of techniques used to calculate rotor power is in order.
ABSTRACT We present results from a suite of 3D high-resolution hydrodynamic simulations of supernova-driven outflows from galactic disc regions with a range of gas surface density, metallicity, and supernova scale height. We use this suite to quantify how outflow properties – particularly the loading factors for mass, metallicity, and energy – vary with these parameters. We find that the winds fall into three broad categories: steady and hot, multiphase and moderately bursty, and cool and highly bursty. The first of these is characterized by efficient metal and energy loading but weak mass loading, the second by moderate loading of mass, metals, and energy, and the third by negligible metal and energy loading but substantial mass loading. The most important factor in determining the kind of wind a galaxy will produce is the ratio of supernova to gas scale heights, with the latter set by a combination of supernova rate, metallicity-dependent cooling rate, and the gravitational potential. These often combine in counterintuitive ways – for example increased cooling causes cold clouds to sink into the galactic midplane more rapidly, lowering the volume-filling factor of dense gas and making the environment more favourable for strong winds. Our findings suggest that the nature of galactic winds is likely highly sensitive to phenomena such as runaway stars occurring at a large height and dense gas and are poorly captured in most simulations, and that metal loading factors for type Ia supernovae may be substantially larger than those for type II, with important implications for galactic chemical evolution.
This paper presents the acausal modeling, validation, and control of floating offshore wind turbines (FOWTs). The model simulates the IEA‐15 MW reference turbine and the semi‐submersible VolturnUS‐S platform utilizing a Control‐oriented, Reconfigurable, and Acausal Floating Turbine Simulator (CRAFTS), which integrates the key coupled aero‐hydro‐elasto‐servo dynamics and is being developed by authors at the University of Central Florida. Verification and validation are conducted using numerical data from the industry‐standard simulation platform OpenFAST and experimental data from the Floating Offshore‐wind and Controls Advanced Laboratory (FOCAL) project, in which the authors were involved. Numerical results demonstrate the model's ability to qualitatively capture loads and responses across various load cases, highlighting the impact of the control system under different wind and wave conditions and opening new opportunities for optimizing FOWT designs. This paper provides wind turbine researchers with valuable insights into system characteristics, system frequencies, damping effects, and internal reaction forces, serving as a reference for future studies in FOWT modeling and control.
Deployable entry vehicles provide an enabling capability for large payloads at destinations across the solar system. The primary advantage of such systems derives from the fundamental limitation of available launch vehicle shroads preventing tried-and-true rigid aeroshell technology from being applied. However, little mission-pull exists for deployable entry vehicle technology for missions where the entry vehicle is the primary payload and can fit within available shrouds. The propensity toward applying rigid aeroshell technology is logical given that deployable concepts are higher risk and provide no marginal scientific benefit for missions where a rigid aeroshell is an option. A movement toward flying smaller spacecraft has emerged in recent years and is revealing novel applications for deployable entry vehicles as secondary payloads. Secondary payloads must minimize interference with the primary payload in order to ease accomodation. Advancements in small spacecraft technology have caught the attention of government entities who are now recognizing the high value proposition of small spacecraft for frequent, incremental technology development and science return rather than the traditional high cost and infrequent big-bang approaches. Worldwide, well over 200 small satellites (mass < 180 kg) have launched in recent years for such purposes as eduation, technology development, science, commerce, and defense. The Adaptable Deployable Entry and Placement Technology (ADEPT), a mechanically-deployable entry vehicle technology, has been under development at NASA since 2011. An Ames Center Innovation Fund (CIF) proposal was recently awarded to study ADEPT for EDL of small payloads. The unique capability of ADEPT for small science payloads comes from its ability to stow within a slender volume and deploy passively to achieve a mass-efficient drag surface with a high heat rate capability. The low ballistic coefficient results in entry heating and mechanical loads that can be met by a revolutionary three-dimensionally woven carbon fabric supported with rib structures. This carbon fabric has test-proven capability as both primary structure and payload thermal protection. This presentation will inform the EDL community of the capability of ADEPT for delivery of small secondary payloads as well as provide status of on-going technology development activities. Results of a mission applications study for secondary payloads to Mars and Venus will be presented. A primary goal of these mission design activities is to integrate high-value verification approaches wherever possible. For example, for some mission applications it is possible to test in NASA arc jet facilities at full-scale. This enables endto- end thermostructural requirement verification with a single test. Another important challenge that needs to be solved is the reliability of the decelerator to achieve the desired shape using simple passive mechanical actuators (such as springs) that do not require power from the primary payload for deployment. A configuration of ADEPT is being prototyped and used to evaluate the functionality and reliability of the deployment mechanism through bench-top testing. The prototype will then be tested in a low speed wind tunnel in late 2014 to evaluate aerodynamic load-carrying capability and deflected shape. Results of the deployment testing and status of wind tunnel test plans will be discussed.
An investigation was undertaken to determine the character and importance of the transition phase between the ground run and steady climb in the takeoff of an airplane and the effects of various factors on this phase and on the airborne part of the takeoff as a whole. The information was obtained from a series of step-by-step integrations, which defined the motion of the airplane during the transition and which were based on data derived from actual takeoff tests of a Verville AT airplane. Both normal and zoom takeoffs under several loading and takeoff speed conditions were considered. The effects of a moderate wind with a corresponding wind gradient and the effect of proximity of the ground were also investigated. The results show that, for normal takeoffs, the best transition was realized at the lowest possible takeoff speed. Moreover, this speed gave the shortest overall takeoff distance for normal takeoffs. Zoom takeoffs required a shorter overall takeoff run than normal takeoffs, particularly with a heavy landing, if the obstacle to be cleared was sufficiently high (greater than 50 feet); no advantage was indicated to the airplane with a light loading if the height to be cleared was less. The error resulting from the neglect of the transition in the calculation of the airborne distance of takeoff was found to vary from 4% with the heaviest loading considered to -4% with the lightest loading for normal takeoffs over a 100-ft obstacle; the percentage error was twice as great for a 50-foot obstacle. For zoom takeoffs the error attained much greater values. The average wind gradient corresponding to a 5-mile-per-hour surface wind reduced the airborne distance required to clear a 50-foot obstacle by about 9% with the lightest loading and 16% with the heaviest loading; for both cases. The overall reduction due to this wind was approximately twice that resulting from the wind gradient alone. A simple expression for the reduction of observed takeoff performance to no-wind conditions is presented. Ground effect is shown to reduce the airborne distance to attain a height of 50 foot by 10% with the lightest loading and 16% with the heaviest loading; for a 100-foot obstacle the percentage reduction was about 1/2 as great.
In order to investigate the impact of ascent element aerodynamics on the flight constraints of the Space Shuttle, a system of programs was developed which allowed the assessment of wind-tunnel data, the extraction of aerodynamic coeffients, the identification of critical structural load indicators and their margins, the calculation of mission-unique shaping parameters and the evaluation of prelaunch wind measurements.
A spacecraft payload flown on a launch vehicle experiences dynamic loads. The dynamic loads are caused by various phenomena ranging from the start-up of the launch vehicle engine to wind gusts. A spacecraft payload should be designed to meet launch vehicle dynamic loads. One of the major steps taken towards determining the dynamic loads is to correlate the finite element model of the spacecraft with the test results of a modal survey test. A test-verified finite element model of the spacecraft should possess the same spatial properties (stiffness, mass, and damping) and modal properties (frequencies and mode shapes) as the test hardware representing the spacecraft. The test-verified and correlated finite element model of the spacecraft is then coupled with the finite element model of the launch vehicle for analysis of loads and stress. Modal survey testing, verification of a finite element model, and modification of the finite element model to match the modal survey test results can easily be accomplished if the spacecraft structure is simple. However, this is rarely the case. A simple structure here is defined as a structure where the influence of nonlinearity between force and displacement (uncertainty in a test, for example, with errors in input and output), and the influence of damping (structural, coulomb, and viscous) are not pronounced. The objective of this study is to develop system identification and correlation methods with the focus on the structural systems that possess nonproportional damping. Two approaches to correct the nonproportional damping matrix of a truss structure were studied, and have been implemented on truss-like structures such as the National Aeronautics and Space Administration's space station truss. The results of this study showed nearly 100 percent improvement of the correlated eigensystem over the analytical eigensystem. The first method showed excellent results with up to three modes used in the system identification process. The second method could handle more modes, but required more computer usage time, and the results were less accurate than those of the first method.
The Control System for the Mod-2 wind turbine system is required to provide not only for startup, RPM regulation, maximizing or regulating power, and stopping the rotor, but also for load limiting, especially in the power train. Early operations with above-rated winds revealed an instability which was caused primarily by coupling between the quill shaft and the rotor air loads. This instability caused the first of several major Mod-2 Control System changes which are reviewed in the paper.
The Simulink-based Simulation Architecture for Evaluating Controls for Aerospace Vehicles (SAREC-ASV) was modified to incorporate linear models representing aeroservoelastic characteristics of the SemiSpan SuperSonic Transport (S4T) wind-tunnel model. The S4T planform is for a Technology Concept Aircraft (TCA) design from the 1990s. The model has three control surfaces and is instrumented with accelerometers and strain gauges. Control laws developed for wind-tunnel testing for Ride Quality Enhancement, Gust Load Alleviation, and Flutter Suppression System functions were implemented in the simulation. The simulation models open- and closed-loop response to turbulence and to control excitation. It provides time histories for closed-loop stable conditions above the open-loop flutter boundary. The simulation is useful for assessing the potential impact of closed-loop control rate and position saturation. It also provides a means to assess fidelity of system identification procedures by providing time histories for a known plant model, with and without unmeasured turbulence as a disturbance. Sets of linear models representing different Mach number and dynamic pressure conditions were implemented as MATLAB Linear Time Invariant (LTI) objects. Configuration changes were implemented by selecting which LTI object to use in a Simulink template block. A limited comparison of simulation versus wind-tunnel results is shown.
For abstract, see volume 1 N77-13130.
For abstract, see volume 1 N77-13130.