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Sample records for aeroacoustic rotor test

  1. Aeroacoustic and Performance Simulations of a Test Scale Open Rotor

    NASA Technical Reports Server (NTRS)

    Claus, Russell W.

    2013-01-01

    This paper explores a comparison between experimental data and numerical simulations of the historical baseline F31/A31 open rotor geometry. The experimental data were obtained at the NASA Glenn Research Center s Aeroacoustic facility and include performance and noise information for a variety of flow speeds (matching take-off and cruise). The numerical simulations provide both performance and aeroacoustic results using the NUMECA s Fine-Turbo analysis code. A non-linear harmonic method is used to capture the rotor/rotor interaction.

  2. Tilt rotor aircraft aeroacoustics

    NASA Technical Reports Server (NTRS)

    George, Albert R.; Smith, Charles A.; Maisel, Martin D.; Brieger, John T.

    1989-01-01

    This paper studies the state of knowledge and the needed improvement in noise methodology and measurements for tilt rotor aircraft. Similarities and differences between tilt rotor aeroacoustic conditions and helicopter and propeller experience are identified. A discussion of the possible principal noise mechanisms throughout the flight envelope shows a need for further experimental and analytical investigations to develop an adequate understanding of the important sources and influencing factors. Existing experimental data from flight tests suggest terminal area noise reduction by operating within certain portions of the conversion flight envelope. Prediction methods are found to provide approximate indications only for low frequency harmonic and broadband noise for several of the tilt rotor's operating conditions. The acoustic effects of the hover case 'fountain' flow are pronounced and need further research. Impulsive noise and high frequency harmonic noise remain problems, as on helicopters, pending major improvements in wake, unsteady aerodynamics, and acoustics methodology.

  3. Tilt rotor hover aeroacoustics

    NASA Technical Reports Server (NTRS)

    Coffen, Charles David

    1992-01-01

    The methodology, results, and conclusions of a study of tilt rotor hover aeroacoustics and aerodynamics are presented. Flow visualization and hot wire velocity measurement were performed on a 1/12-scale model of the XV-15 Tilt Rotor Aircraft in hover. The wing and fuselage below the rotor cause a complex recirculating flow. Results indicate the physical dimensions and details of the flow including the relative unsteadiness and turbulence characteristics of the flow. Discrete frequency harmonic thickness and the loading noise mechanism were predicted using WOPWOP for the standard metal blades and the Advanced Technology Blades. The recirculating flow created by the wing below the rotor is a primary sound mechanism for a hovering tilt rotor. The effects of dynamic blade response should be included for fountain flow conditions which produce impulsive blade loading. Broadband noise mechanisms were studied using Amiet's method with azimuthally varying turbulence characteristics derived from the measurements. The recirculating fountain flow with high turbulence levels in the recirculating zone is the dominant source of broadband noise for a hovering rotor. It is shown that tilt rotor hover aeroacoustic noise mechanisms are now understood. Noise predictions can be made based on reasonably accurate aerodynamic models developed here.

  4. Open Rotor Aeroacoustic Modelling

    NASA Technical Reports Server (NTRS)

    Envia, Edmane

    2012-01-01

    Owing to their inherent fuel efficiency, there is renewed interest in developing open rotor propulsion systems that are both efficient and quiet. The major contributor to the overall noise of an open rotor system is the propulsor noise, which is produced as a result of the interaction of the airstream with the counter-rotating blades. As such, robust aeroacoustic prediction methods are an essential ingredient in any approach to designing low-noise open rotor systems. To that end, an effort has been underway at NASA to assess current open rotor noise prediction tools and develop new capabilities. Under this effort, high-fidelity aerodynamic simulations of a benchmark open rotor blade set were carried out and used to make noise predictions via existing NASA open rotor noise prediction codes. The results have been compared with the aerodynamic and acoustic data that were acquired for this benchmark open rotor blade set. The emphasis of this paper is on providing a summary of recent results from a NASA Glenn effort to validate an in-house open noise prediction code called LINPROP which is based on a high-blade-count asymptotic approximation to the Ffowcs-Williams Hawkings Equation. The results suggest that while predicting the absolute levels may be difficult, the noise trends are reasonably well predicted by this approach.

  5. Open Rotor Aeroacoustic Modeling

    NASA Technical Reports Server (NTRS)

    Envia, Edmane

    2012-01-01

    Owing to their inherent fuel efficiency, there is renewed interest in developing open rotor propulsion systems that are both efficient and quiet. The major contributor to the overall noise of an open rotor system is the propulsor noise, which is produced as a result of the interaction of the airstream with the counter-rotating blades. As such, robust aeroacoustic prediction methods are an essential ingredient in any approach to designing low-noise open rotor systems. To that end, an effort has been underway at NASA to assess current open rotor noise prediction tools and develop new capabilities. Under this effort, high-fidelity aerodynamic simulations of a benchmark open rotor blade set were carried out and used to make noise predictions via existing NASA open rotor noise prediction codes. The results have been compared with the aerodynamic and acoustic data that were acquired for this benchmark open rotor blade set. The emphasis of this paper is on providing a summary of recent results from a NASA Glenn effort to validate an in-house open noise prediction code called LINPROP which is based on a high-blade-count asymptotic approximation to the Ffowcs-Williams Hawkings Equation. The results suggest that while predicting the absolute levels may be difficult, the noise trends are reasonably well predicted by this approach.

  6. Full Scale Rotor Aeroacoustic Predictions and the Link to Model Scale Rotor Data

    NASA Technical Reports Server (NTRS)

    Boyd, D. Douglas, Jr.; Burley, Casey L.; Conner, David A.

    2004-01-01

    The NASA Aeroacoustic Prediction System (NAPS) is used to establish a link between model-scale and full-scale rotor predictions and is partially validated against measured wind tunnel and flight aeroacoustic data. The prediction approach of NAPS couples a comprehensive rotorcraft analysis with acoustic source noise and propagation codes. The comprehensive analysis selected for this study is CAMRAD-II, which provides the performance/trim/wake solution for a given rotor or flight condition. The post-trim capabilities of CAMRAD-II are used to compute high-resolution sectional airloads for the acoustic tone noise analysis, WOPMOD. The tone noise is propagated to observers on the ground with the propagation code, RNM (Rotor Noise Model). Aeroacoustic predictions are made with NAPS for an isolated rotor and compared to results of the second Harmonic Aeroacoustic Rotor Test (HART-II) program, which tested a 40% dynamically and Mach-scaled BO-105 main rotor at the DNW. The NAPS is validated with comparisons for three rotor conditions: a baseline condition and two Higher Harmonic Control (HHC) conditions. To establish a link between model and full-scale rotor predictions, a full-scale BO-105 main rotor input deck for NAPS is created from the 40% scale rotor input deck. The full-scale isolated rotor predictions are then compared to the model predictions. The comparisons include aerodynamic loading, acoustic levels, and acoustic pressure time histories for each of the three conditions. With this link established, full-scale predictions are made for a range of descent flight conditions and compared with measured trends from the recent Rotorcraft Operational Noise Abatement Procedures (RONAP) flight test conducted by DLR and ONERA. Additionally, the effectiveness of two HHC conditions from the HART-II program is demonstrated for the full-scale rotor in flight.

  7. Open Rotor Aeroacoustic Installation Effects for Conventional and Unconventional Airframes

    NASA Technical Reports Server (NTRS)

    Czech, Michael J.; Thomas, Russell H.

    2013-01-01

    As extensive experimental campaign was performed to study the aeroacoustic installation effects of an open rotor with respect to both a conventional tube and wing type airframe and an unconventional hybrid wing body airframe. The open rotor rig had two counter rotating rows of blades each with eight blades of a design originally flight tested in the 1980s. The aeroacoustic installation effects measured in an aeroacoustic wind tunnel included those from flow effects due to inflow distortion or wake interaction and acoustic propagation effects such as shielding and reflection. The objective of the test campaign was to quantify the installation effects for a wide range of parameters and configurations derived from the two airframe types. For the conventional airframe, the open rotor was positioned in increments in front of and then over the main wing and then in positions representative of tail mounted aircraft with a conventional tail, a T-tail and a U-tail. The interaction of the wake of the open rotor as well as acoustic scattering results in an increase of about 10 dB when the rotor is positioned in front of the main wing. When positioned over the main wing a substantial amount of noise reduction is obtained and this is also observed for tail-mounted installations with a large U-tail. For the hybrid wing body airframe, the open rotor was positioned over the airframe along the centerline as well as off-center representing a twin engine location. A primary result was the documentation of the noise reduction from shielding as a function of the location of the open rotor upstream of the trailing edge of the hybrid wing body. The effects from vertical surfaces and elevon deflection were also measured. Acoustic lining was specially designed and inserted flush with the elevon and airframe surface, the result was an additional reduction in open rotor noise propagating to the far field microphones. Even with the older blade design used, the experiment provided

  8. Aeroacoustics and aerodynamic performance of a rotor with flatback airfoils.

    SciTech Connect

    Paquette, Joshua A.; Barone, Matthew Franklin; Christiansen, Monica; Simley, Eric

    2010-06-01

    The aerodynamic performance and aeroacoustic noise sources of a rotor employing flatback airfoils have been studied in field test campaign and companion modeling effort. The field test measurements of a sub-scale rotor employing nine meter blades include both performance measurements and acoustic measurements. The acoustic measurements are obtained using a 45 microphone beamforming array, enabling identification of both noise source amplitude and position. Semi-empirical models of flatback airfoil blunt trailing edge noise are developed and calibrated using available aeroacoustic wind tunnel test data. The model results and measurements indicate that flatback airfoil noise is less than drive train noise for the current test turbine. It is also demonstrated that the commonly used Brooks, Pope, and Marcolini model for blunt trailing edge noise may be over-conservative in predicting flatback airfoil noise for wind turbine applications.

  9. Calculation of the Aerodynamic Behavior of the Tilt Rotor Aeroacoustic Model (TRAM) in the DNW

    NASA Technical Reports Server (NTRS)

    Johnson, Wayne

    2001-01-01

    Comparisons of measured and calculated aerodynamic behavior of a tiltrotor model are presented. The test of the Tilt Rotor Aeroacoustic Model (TRAM) with a single, 1/4-scale V- 22 rotor in the German-Dutch Wind Tunnel (DNW) provides an extensive set of aeroacoustic, performance, and structural loads data. The calculations were performed using the rotorcraft comprehensive analysis CAMRAD II. Presented are comparisons of measured and calculated performance and airloads for helicopter mode operation, as well as calculated induced and profile power. An aerodynamic and wake model and calculation procedure that reflects the unique geometry and phenomena of tiltrotors has been developed. There are major differences between this model and the corresponding aerodynamic and wake model that has been established for helicopter rotors. In general, good correlation between measured and calculated performance and airloads behavior has been shown. Two aspects of the analysis that clearly need improvement are the stall delay model and the trailed vortex formation model.

  10. Open Rotor Computational Aeroacoustic Analysis with an Immersed Boundary Method

    NASA Technical Reports Server (NTRS)

    Brehm, Christoph; Barad, Michael F.; Kiris, Cetin C.

    2016-01-01

    Reliable noise prediction capabilities are essential to enable novel fuel efficient open rotor designs that can meet the community and cabin noise standards. Toward this end, immersed boundary methods have reached a level of maturity so that they are being frequently employed for specific real world applications within NASA. This paper demonstrates that our higher-order immersed boundary method provides the ability for aeroacoustic analysis of wake-dominated flow fields generated by highly complex geometries. This is the first of a kind aeroacoustic simulation of an open rotor propulsion system employing an immersed boundary method. In addition to discussing the peculiarities of applying the immersed boundary method to this moving boundary problem, we will provide a detailed aeroacoustic analysis of the noise generation mechanisms encountered in the open rotor flow. The simulation data is compared to available experimental data and other computational results employing more conventional CFD methods. The noise generation mechanisms are analyzed employing spectral analysis, proper orthogonal decomposition and the causality method.

  11. Open Rotor Computational Aeroacoustic Analysis with an Immersed Boundary Method

    NASA Technical Reports Server (NTRS)

    Brehm, Christoph; Barad, Michael F.; Kiris, Cetin C.

    2016-01-01

    Reliable noise prediction capabilities are essential to enable novel fuel efficient open rotor designs that can meet the community and cabin noise standards. Toward this end, immersed boundary methods have reached a level of maturity where more and more complex flow problems can be tackled with this approach. This paper demonstrates that our higher-order immersed boundary method provides the ability for aeroacoustic analysis of wake-dominated flow fields generated by a contra-rotating open rotor. This is the first of a kind aeroacoustic simulation of an open rotor propulsion system employing an immersed boundary method. In addition to discussing the methodologies of how to apply the immersed boundary method to this moving boundary problem, we will provide a detailed validation of the aeroacoustic analysis approach employing the Launch Ascent and Vehicle Aerodynamics (LAVA) solver. Two free-stream Mach numbers with M=0.2 and M=0.78 are considered in this analysis that are based on the nominally take-off and cruise flow conditions. The simulation data is compared to available experimental data and other computational results employing more conventional CFD methods. Spectral analysis is used to determine the dominant wave propagation pattern in the acoustic near-field.

  12. Computational analysis of high resolution unsteady airloads for rotor aeroacoustics

    NASA Technical Reports Server (NTRS)

    Quackenbush, Todd R.; Lam, C.-M. Gordon; Wachspress, Daniel A.; Bliss, Donald B.

    1994-01-01

    The study of helicopter aerodynamic loading for acoustics applications requires the application of efficient yet accurate simulations of the velocity field induced by the rotor's vortex wake. This report summarizes work to date on the development of such an analysis, which builds on the Constant Vorticity Contour (CVC) free wake model, previously implemented for the study of vibratory loading in the RotorCRAFT computer code. The present effort has focused on implementation of an airload reconstruction approach that computes high resolution airload solutions of rotor/rotor-wake interactions required for acoustics computations. Supplementary efforts on the development of improved vortex core modeling, unsteady aerodynamic effects, higher spatial resolution of rotor loading, and fast vortex wake implementations have substantially enhanced the capabilities of the resulting software, denoted RotorCRAFT/AA (AeroAcoustics). Results of validation calculations using recently acquired model rotor data show that by employing airload reconstruction it is possible to apply the CVC wake analysis with temporal and spatial resolution suitable for acoustics applications while reducing the computation time required by one to two orders of magnitude relative to that required by direct calculations. Promising correlation with this body of airload and noise data has been obtained for a variety of rotor configurations and operating conditions.

  13. Aeroacoustic Codes For Rotor Harmonic and BVI Noise--CAMRAD.Mod1/HIRES

    NASA Technical Reports Server (NTRS)

    Brooks, Thomas F.; Boyd, D. Douglas, Jr.; Burley, Casey L.; Jolly, J. Ralph, Jr.

    1996-01-01

    This paper presents a status of non-CFD aeroacoustic codes at NASA Langley Research Center for the prediction of helicopter harmonic and Blade-Vortex Interaction (BVI) noise. The prediction approach incorporates three primary components: CAMRAD.Mod1 - a substantially modified version of the performance/trim/wake code CAMRAD; HIRES - a high resolution blade loads post-processor; and WOPWOP - an acoustic code. The functional capabilities and physical modeling in CAMRAD.Mod1/HIRES will be summarized and illustrated. A new multi-core roll-up wake modeling approach is introduced and validated. Predictions of rotor wake and radiated noise are compared with to the results of the HART program, a model BO-105 windtunnel test at the DNW in Europe. Additional comparisons are made to results from a DNW test of a contemporary design four-bladed rotor, as well as from a Langley test of a single proprotor (tiltrotor) three-bladed model configuration. Because the method is shown to help eliminate the necessity of guesswork in setting code parameters between different rotor configurations, it should prove useful as a rotor noise design tool.

  14. Tests of anechoic chamber for aeroacoustics investigations

    NASA Astrophysics Data System (ADS)

    Palchikovskiy, V. V.; Bersenev, Yu. V.; Makashov, S. Yu.; Belyaev, I. V.; Korin, I. A.; Sorokin, E. V.; Khramtsov, I. V.; Kustov, O. Yu.

    2016-10-01

    The paper presents the results of qualification tests in the new anechoic chamber of Perm National Research Polytechnic University (PNRPU) built in 2014-2015 and evaluation of the chamber quality in aeroacoustic experiments. It describes design features of the chamber and its sound-absorption lining. The qualification tests were carried out with tonal and broadband noise sources in the frequency range 100 Hz - 20 kHz for two different cases of the source arrangement. In every case, measurements were performed in three directions by traverse microphones. Qualification tests have determined that in the chamber there is a free acoustic field within radius of 2 m for tonal noise and 3 m for broadband noise. There was also evaluated acoustic quality of the chamber by measurements of the jet noise and vortex ring noise. The results of the experiments demonstrate that PNRPU anechoic chamber allows the aeroacoustic measurements to be performed to obtain quantitative results.

  15. Prediction of the Aero-Acoustic Performance of Open Rotors

    NASA Technical Reports Server (NTRS)

    VanZante, Dale; Envia, Edmane

    2014-01-01

    The rising cost of jet fuel has renewed interest in contrarotating open rotor propulsion systems. Contemporary design methods offer the potential to maintain the inherently high aerodynamic efficiency of open rotors while greatly reducing their noise output, something that was not feasible in the 1980's designs. The primary source mechanisms of open rotor noise generation are thought to be the front rotor wake and tip vortex interacting with the aft rotor. In this paper, advanced measurement techniques and high-fidelity prediction tools are used to gain insight into the relative importance of the contributions to the open rotor noise signature of the front rotor wake and rotor tip vortex. The measurements include three-dimensional particle image velocimetry of the intra-rotor flowfield and the acoustic field of a model-scale open rotor. The predictions provide the unsteady flowfield and the associated acoustic field. The results suggest that while the front rotor tip vortex can have a significant influence on the blade passing tone noise produced by the aft rotor, the front rotor wake plays the decisive role in the generation of the interaction noise produced as a result of the unsteady aerodynamic interaction of the two rotors. At operating conditions typical of takeoff and landing operations, the interaction noise level is easily on par with that generated by the individual rotors, and in some cases is even higher. This suggests that a comprehensive approach to reducing open rotor noise should include techniques for mitigating the wake of the front rotor as well as eliminating the interaction of the front rotor tip vortex with the aft rotor blade tip.

  16. Aeromechanics and Aeroacoustics Predictions of the Boeing-SMART Rotor Using Coupled-CFD/CSD Analyses

    NASA Technical Reports Server (NTRS)

    Bain, Jeremy; Sim, Ben W.; Sankar, Lakshmi; Brentner, Ken

    2010-01-01

    This paper will highlight helicopter aeromechanics and aeroacoustics prediction capabilities developed by Georgia Institute of Technology, the Pennsylvania State University, and Northern Arizona University under the Helicopter Quieting Program (HQP) sponsored by the Tactical Technology Office of the Defense Advanced Research Projects Agency (DARPA). First initiated in 2004, the goal of the HQP was to develop high fidelity, state-of-the-art computational tools for designing advanced helicopter rotors with reduced acoustic perceptibility and enhanced performance. A critical step towards achieving this objective is the development of rotorcraft prediction codes capable of assessing a wide range of helicopter configurations and operations for future rotorcraft designs. This includes novel next-generation rotor systems that incorporate innovative passive and/or active elements to meet future challenging military performance and survivability goals.

  17. Tiltrotor Aeroacoustic Code (TRAC) Prediction Assessment and Initial Comparisons with Tram Test Data

    NASA Technical Reports Server (NTRS)

    Burley, Casey L.; Brooks, Thomas F.; Charles, Bruce D.; McCluer, Megan

    1999-01-01

    A prediction sensitivity assessment to inputs and blade modeling is presented for the TiltRotor Aeroacoustic Code (TRAC). For this study, the non-CFD prediction system option in TRAC is used. Here, the comprehensive rotorcraft code, CAMRAD.Mod1, coupled with the high-resolution sectional loads code HIRES, predicts unsteady blade loads to be used in the noise prediction code WOPWOP. The sensitivity of the predicted blade motions, blade airloads, wake geometry, and acoustics is examined with respect to rotor rpm, blade twist and chord, and to blade dynamic modeling. To accomplish this assessment, an interim input-deck for the TRAM test model and an input-deck for a reference test model are utilized in both rigid and elastic modes. Both of these test models are regarded as near scale models of the V-22 proprotor (tiltrotor). With basic TRAC sensitivities established, initial TRAC predictions are compared to results of an extensive test of an isolated model proprotor. The test was that of the TiltRotor Aeroacoustic Model (TRAM) conducted in the Duits-Nederlandse Windtunnel (DNW). Predictions are compared to measured noise for the proprotor operating over an extensive range of conditions. The variation of predictions demonstrates the great care that must be taken in defining the blade motion. However, even with this variability, the predictions using the different blade modeling successfully capture (bracket) the levels and trends of the noise for conditions ranging from descent to ascent.

  18. Open Rotor Spin Test

    NASA Video Gallery

    An open rotor, also known as a high-speed propeller, is tested in a wind tunnel. The propeller moves much more quickly than a standard propeller, and the blades of the propeller are shaped differen...

  19. Arrays of Miniature Microphones for Aeroacoustic Testing

    NASA Technical Reports Server (NTRS)

    Shams, Qamar A.; Humphreys, William M.; Sealey, Bradley S.; Bartram, Scott M.; Zuckewar, Allan J.; Comeaux, Toby; Adams, James K.

    2007-01-01

    A phased-array system comprised of custom-made and commercially available microelectromechanical system (MEMS) silicon microphones and custom ancillary hardware has been developed for use in aeroacoustic testing in hard-walled and acoustically treated wind tunnels. Recent advances in the areas of multi-channel signal processing and beam forming have driven the construction of phased arrays containing ever-greater numbers of microphones. Traditional obstacles to this trend have been posed by (1) the high costs of conventional condenser microphones, associated cabling, and support electronics and (2) the difficulty of mounting conventional microphones in the precise locations required for high-density arrays. The present development overcomes these obstacles. One of the hallmarks of the new system is a series of fabricated platforms on which multiple microphones can be mounted. These mounting platforms, consisting of flexible polyimide circuit-board material (see left side of figure), include all the necessary microphone power and signal interconnects. A single bus line connects all microphones to a common power supply, while the signal lines terminate in one or more data buses on the sides of the circuit board. To minimize cross talk between array channels, ground lines are interposed as shields between all the data bus signal lines. The MEMS microphones are electrically connected to the boards via solder pads that are built into the printed wiring. These flexible circuit boards share many characteristics with their traditional rigid counterparts, but can be manufactured much thinner, as small as 0.1 millimeter, and much lighter with boards weighing as much as 75 percent less than traditional rigid ones. For a typical hard-walled wind-tunnel installation, the flexible printed-circuit board is bonded to the tunnel wall and covered with a face sheet that contains precise cutouts for the microphones. Once the face sheet is mounted, a smooth surface is established over

  20. A comparison of the acoustic and aerodynamic measurements of a model rotor tested in two anechoic wind tunnels

    NASA Technical Reports Server (NTRS)

    Boxwell, D. A.; Schmitz, F. H.; Splettstoesser, W. R.; Schultz, K. J.; Lewy, S.; Caplot, M.

    1986-01-01

    Two aeroacoustic facilities--the CEPRA 19 in France and the DNW in the Netherlands--are compared. The two facilities have unique acoustic characteristics that make them appropriate for acoustic testing of model-scale helicopter rotors. An identical pressure-instrumented model-scale rotor was tested in each facility and acoustic test results are compared with full-scale-rotor test results. Blade surface pressures measured in both tunnels were used to correlated nominal rotor operating conditions in each tunnel, and also used to assess the steadiness of the rotor in each tunnel's flow. In-the-flow rotor acoustic signatures at moderate forward speeds (35-50 m/sec) are presented for each facility and discussed in relation to the differences in tunnel geometries and aeroacoustic characteristics. Both reports are presented in appendices to this paper. ;.);

  1. Aerodynamic and acoustic test of a United Technologies model scale rotor at DNW

    NASA Technical Reports Server (NTRS)

    Yu, Yung H.; Liu, Sandy R.; Jordan, Dave E.; Landgrebe, Anton J.; Lorber, Peter F.; Pollack, Michael J.; Martin, Ruth M.

    1990-01-01

    The UTC model scale rotors, the DNW wind tunnel, the AFDD rotary wing test stand, the UTRC and AFDD aerodynamic and acoustic data acquisition systems, and the scope of test matrices are discussed and an introduction to the test results is provided. It is pointed out that a comprehensive aero/acoustic database of several configurations of the UTC scaled model rotor has been created. The data is expected to improve understanding of rotor aerodynamics, acoustics, and dynamics, and lead to enhanced analytical methodology and design capabilities for the next generation of rotorcraft.

  2. Wind Tunnel Test of the SMART Active Flap Rotor

    NASA Technical Reports Server (NTRS)

    Straub, Friedrich K.; Anand, Vaidyanthan R.; Birchette, Terrence S.; Lau, Benton H.

    2009-01-01

    Boeing and a team from Air Force, NASA, Army, DARPA, MIT, UCLA, and U. of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. The Boeing SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing edge flap on each blade. The eleven-week test program evaluated the forward flight characteristics of the active-flap rotor at speeds up to 155 knots, gathered data to validate state-of-the-art codes for rotor aero-acoustic analysis, and quantified the effects of open and closed loop active flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness of the active flap control on noise and vibration was conclusively demonstrated. Results showed significant reductions up to 6dB in blade-vortex-interaction and in-plane noise, as well as reductions in vibratory hub loads up to 80%. Trailing-edge flap deflections were controlled within 0.1 degrees of the commanded value. The impact of the active flap on control power, rotor smoothing, and performance was also demonstrated. Finally, the reliability of the flap actuation system was successfully proven in more than 60 hours of wind-tunnel testing.

  3. Acoustic impedance testing for aeroacoustic applications

    NASA Astrophysics Data System (ADS)

    Schultz, Todd

    Accurate acoustic propagation models are required to characterize and subsequently reduce aircraft engine noise. These models ultimately rely on acoustic impedance measurements of candidate materials used in sound-absorbing liners. The standard two-microphone method (TMM) is widely used to estimate acoustic impedance but is limited in frequency range and does not provide uncertainty estimates, which are essential for data quality assessment and model validation. This dissertation presents a systematic framework to estimate uncertainty and extend the frequency range of acoustic impedance testing. Uncertainty estimation for acoustic impedance data using the TMM is made via two methods. The first employs a standard analytical technique based on linear perturbations and provides useful scaling information. The second uses a Monte Carlo technique that permits the propagation of arbitrarily large uncertainties. Both methods are applied to the TMM for simulated data representative of sound-hard and sound-soft acoustic materials. The results indicate that the analytical technique can lead to false conclusions about the magnitude and importance of specific error sources. Furthermore, the uncertainty in acoustic impedance is strongly dependent on the frequency and the uncertainty in the microphone locations. Next, an increased frequency range of acoustic impedance testing is investigated via two methods. The first method reduces the size of the test specimen (from 25.4 mm square to 8.5 mm square) and uses the standard TMM. This method has issues concerning specimen nonuniformity because the small specimens may not be representative of the material. The second method increases the duct cross section and, hence, the required complexity of the sound field propagation model. A comparison among all three methods is conducted for each of the three specimens: two different ceramic tubular specimens and a single degree-of-freedom liner. The results show good agreement between the

  4. A Comparison of Computational Aeroacoustic Prediction Methods for Transonic Rotor Noise

    NASA Technical Reports Server (NTRS)

    Brentner, Kenneth S.; Lyrintzis, Anastasios; Koutsavdis, Evangelos K.

    1996-01-01

    This paper compares two methods for predicting transonic rotor noise for helicopters in hover and forward flight. Both methods rely on a computational fluid dynamics (CFD) solution as input to predict the acoustic near and far fields. For this work, the same full-potential rotor code has been used to compute the CFD solution for both acoustic methods. The first method employs the acoustic analogy as embodied in the Ffowcs Williams-Hawkings (FW-H) equation, including the quadrupole term. The second method uses a rotating Kirchhoff formulation. Computed results from both methods are compared with one other and with experimental data for both hover and advancing rotor cases. The results are quite good for all cases tested. The sensitivity of both methods to CFD grid resolution and to the choice of the integration surface/volume is investigated. The computational requirements of both methods are comparable; in both cases these requirements are much less than the requirements for the CFD solution.

  5. NASA Hybrid Wing Aircraft Aeroacoustic Test Documentation Report

    NASA Technical Reports Server (NTRS)

    Heath, Stephanie L.; Brooks, Thomas F.; Hutcheson, Florence V.; Doty, Michael J.; Bahr, Christopher J.; Hoad, Danny; Becker, Lawrence; Humphreys, William M.; Burley, Casey L.; Stead, Dan; Pope, Dennis S.; Spalt, Taylor B.; Kuchta, Dennis H.; Plassman, Gerald E.; Moen, Jaye A.

    2016-01-01

    This report summarizes results of the Hybrid Wing Body (HWB) N2A-EXTE model aeroacoustic test. The N2A-EXTE model was tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel (14x22 Tunnel) from September 12, 2012 until January 28, 2013 and was designated as test T598. This document contains the following main sections: Section 1 - Introduction, Section 2 - Main Personnel, Section 3 - Test Equipment, Section 4 - Data Acquisition Systems, Section 5 - Instrumentation and Calibration, Section 6 - Test Matrix, Section 7 - Data Processing, and Section 8 - Summary. Due to the amount of material to be documented, this HWB test documentation report does not cover analysis of acquired data, which is to be presented separately by the principal investigators. Also, no attempt was made to include preliminary risk reduction tests (such as Broadband Engine Noise Simulator and Compact Jet Engine Simulator characterization tests, shielding measurement technique studies, and speaker calibration method studies), which were performed in support of this HWB test. Separate reports containing these preliminary tests are referenced where applicable.

  6. Acoustic Performance of an Advanced Model Turbofan in Three Aeroacoustic Test Facilities

    NASA Technical Reports Server (NTRS)

    Woodward, Richard P.; Hughes, Christopher E.

    2012-01-01

    A model advanced turbofan was acoustically tested in the NASA Glenn 9- by 15-Foot-Low-Speed Wind Tunnel (LSWT), and in two other aeroacoustic facilities. The Universal Propulsion Simulator (UPS) fan was designed and manufactured by the General Electric Aircraft Engines (GEAE) Company, and featured active core, as well as bypass, flow paths. The reference test configurations were with the metal, M4, rotor with hardwall and treated bypass flow ducts. The UPS fan was tested within an airflow at a Mach number of 0.20 (limited flow data were also acquired at a Mach number of 0.25) which is representative of aircraft takeoff and approach conditions. Comparisons were made between data acquired within the airflow (9x15 LSWT and German-Dutch Wind Tunnel (DNW)) and outside of a free jet (Boeing Low Speed Aero acoustic Facility (LSAF) and DNW). Sideline data were acquired on an 89-in. (nominal 4 fan diameters) sideline using the same microphone assembly and holder in the 9x15 LSWT and DNW facilities. These data showed good agreement for similar UPS operating conditions and configurations. Distortion of fan spectra tonal content through a free jet shear layer was documented, suggesting that in-flow acoustic measurements are required for comprehensive fan noise diagnostics. However, there was good agreement for overall sound power level (PWL) fan noise measurements made both within and outside of the test facility airflow.

  7. Aerodynamic and Aeroacoustic Wind Tunnel Testing of the Orion Spacecraft

    NASA Technical Reports Server (NTRS)

    Ross, James C.

    2011-01-01

    The Orion aerodynamic testing team has completed more than 40 tests as part of developing the aerodynamic and loads databases for the vehicle. These databases are key to achieving good mechanical design for the vehicle and to ensure controllable flight during all potential atmospheric phases of a mission, including launch aborts. A wide variety of wind tunnels have been used by the team to document not only the aerodynamics but the aeroacoustic environment that the Orion might experience both during nominal ascents and launch aborts. During potential abort scenarios the effects of the various rocket motor plumes on the vehicle must be accurately understood. The Abort Motor (AM) is a high-thrust, short duration motor that rapidly separates Orion from its launch vehicle. The Attitude Control Motor (ACM), located in the nose of the Orion Launch Abort Vehicle, is used for control during a potential abort. The 8 plumes from the ACM interact in a nonlinear manner with the four AM plumes which required a carefully controlled test to define the interactions and their effect on the control authority provided by the ACM. Techniques for measuring dynamic stability and for simulating rocket plume aerodynamics and acoustics were improved or developed in the course of building the aerodynamic and loads databases for Orion.

  8. FJ44 Turbofan Engine Test at NASA Glenn Research Center's Aero-Acoustic Propulsion Laboratory

    NASA Technical Reports Server (NTRS)

    Lauer, Joel T.; McAllister, Joseph; Loew, Raymond A.; Sutliff, Daniel L.; Harley, Thomas C.

    2009-01-01

    A Williams International FJ44-3A 3000-lb thrust class turbofan engine was tested in the NASA Glenn Research Center s Aero-Acoustic Propulsion Laboratory. This report presents the test set-up and documents the test conditions. Farfield directivity, in-duct unsteady pressures, duct mode data, and phased-array data were taken and are reported separately.

  9. Prediction of XV-15 tilt rotor discrete frequency aeroacoustic noise with WOPWOP

    NASA Technical Reports Server (NTRS)

    Coffen, Charles D.; George, Albert R.

    1990-01-01

    The results, methodology, and conclusions of noise prediction calculations carried out to study several possible discrete frequency harmonic noise mechanisms of the XV-15 Tilt Rotor Aircraft in hover and helicopter mode forward flight are presented. The mechanisms studied were thickness and loading noise. In particular, the loading noise caused by flow separation and the fountain/ground plane effect were predicted with calculations made using WOPWOP, a noise prediction program developed by NASA Langley. The methodology was to model the geometry and aerodynamics of the XV-15 rotor blades in hover and steady level flight and then create corresponding FORTRAN subroutines which were used an input for WOPWOP. The models are described and the simplifying assumptions made in creating them are evaluated, and the results of the computations are presented. The computations lead to the following conclusions: The fountain/ground plane effect is an important source of aerodynamic noise for the XV-15 in hover. Unsteady flow separation from the airfoil passing through the fountain at high angles of attack significantly affects the predicted sound spectra and may be an important noise mechanism for the XV-15 in hover mode. The various models developed did not predict the sound spectra in helicopter forward flight. The experimental spectra indicate the presence of blade vortex interactions which were not modeled in these calculations. A need for further study and development of more accurate aerodynamic models, including unsteady stall in hover and blade vortex interactions in forward flight.

  10. A Large Hemi-Anechoic Enclosure for Community-Compatible Aeroacoustic Testing of Aircraft Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Cooper, Beth A.

    1993-01-01

    A large hemi-anechoic (absorptive walls and acoustically hard floor) noise control enclosure has been erected around a complex of test stands at the NASA Lewis Research Center in Cleveland, Ohio. This new state-of-the-art Aeroacoustic Propulsion Laboratory (APL) provides an all-weather, semisecure test environment while limiting noise to acceptable levels in surrounding residential neighborhoods. The 39.6 m (130 ft) diameter geodesic dome structure houses the new Nozzle Aeroacoustic Test Rig (NATR), an ejector-powered M = 0.3 free jet facility for acoustic testing of supersonic aircraft exhaust nozzles and turbomachinery. A multi-axis, force-measuring Powered Lift Facility (PLF) stand for testing of Short Takeoff Vertical Landing (STOVL) vehicles is also located within the dome. The design of the Aeroacoustic Propulsion Laboratory efficiently accomodates the research functions of two separate test rigs, one of which (NATR) requires a specialized environment for taking acoustic measurements. Absorptive fiberglass wedge treatment on the interior surface of the dome provides a hemi-anechoic interior environment for obtaining the accurate acoustic measurements required to meet research program goals. The APL is the first known geodesic dome structure to incorporate transmission-loss properties as well as interior absorption into a free-standing, community-compatible, hemi-anechoic test facility.

  11. Aeroacoustic Testing of Wind Turbine Airfoils: February 20, 2004 - February 19, 2008

    SciTech Connect

    Devenport, W.; Burdisso, R. A.; Camargo, H.; Crede, E.; Remillieux, M.; Rasnick, M.; Van Seeters, P.

    2010-05-01

    The U.S. Department of Energy (DOE), working through its National Renewable Energy Laboratory (NREL), is engaged in a comprehensive research effort to improve the understanding of wind turbine aeroacoustics. The motivation for this effort is the desire to exploit the large expanse of low wind speed sites that tend to be close to U.S. load centers. Quiet wind turbines are an inducement to widespread deployment, so the goal of NREL's aeroacoustic research is to develop tools that the U.S. wind industry can use in developing and deploying highly efficient, quiet wind turbines at low wind speed sites. NREL's National Wind Technology Center (NWTC) is implementing a multifaceted approach that includes wind tunnel tests, field tests, and theoretical analyses in direct support of low wind speed turbine development by its industry partners. NWTC researchers are working hand in hand with engineers in industry to ensure that research findings are available to support ongoing design decisions.

  12. Advanced Background Subtraction Applied to Aeroacoustic Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Bahr, Christopher J.; Horne, William C.

    2015-01-01

    An advanced form of background subtraction is presented and applied to aeroacoustic wind tunnel data. A variant of this method has seen use in other fields such as climatology and medical imaging. The technique, based on an eigenvalue decomposition of the background noise cross-spectral matrix, is robust against situations where isolated background auto-spectral levels are measured to be higher than levels of combined source and background signals. It also provides an alternate estimate of the cross-spectrum, which previously might have poor definition for low signal-to-noise ratio measurements. Simulated results indicate similar performance to conventional background subtraction when the subtracted spectra are weaker than the true contaminating background levels. Superior performance is observed when the subtracted spectra are stronger than the true contaminating background levels. Experimental results show limited success in recovering signal behavior for data where conventional background subtraction fails. They also demonstrate the new subtraction technique's ability to maintain a proper coherence relationship in the modified cross-spectral matrix. Beam-forming and de-convolution results indicate the method can successfully separate sources. Results also show a reduced need for the use of diagonal removal in phased array processing, at least for the limited data sets considered.

  13. Research investigation of helicopter main rotor/tail rotor interaction noise

    NASA Technical Reports Server (NTRS)

    Fitzgerald, J.; Kohlhepp, F.

    1988-01-01

    Acoustic measurements were obtained in a Langley 14 x 22 foot Subsonic Wind Tunnel to study the aeroacoustic interaction of 1/5th scale main rotor, tail rotor, and fuselage models. An extensive aeroacoustic data base was acquired for main rotor, tail rotor, fuselage aerodynamic interaction for moderate forward speed flight conditions. The details of the rotor models, experimental design and procedure, aerodynamic and acoustic data acquisition and reduction are presented. The model was initially operated in trim for selected fuselage angle of attack, main rotor tip-path-plane angle, and main rotor thrust combinations. The effects of repositioning the tail rotor in the main rotor wake and the corresponding tail rotor countertorque requirements were determined. Each rotor was subsequently tested in isolation at the thrust and angle of attack combinations for trim. The acoustic data indicated that the noise was primarily dominated by the main rotor, especially for moderate speed main rotor blade-vortex interaction conditions. The tail rotor noise increased when the main rotor was removed indicating that tail rotor inflow was improved with the main rotor present.

  14. Overview of the Space Launch System Ascent Aeroacoustic Environment Test Program

    NASA Technical Reports Server (NTRS)

    Herron, Andrew J.; Crosby, William A.; Reed, Darren K.

    2016-01-01

    Characterization of accurate flight vehicle unsteady aerodynamics is critical for component and secondary structure vibroacoustic design. The Aerosciences Branch at the National Aeronautics and Space Administration (NASA) Marshall Space Flight Center has conducted a test at the NASA Ames Research Center (ARC) Unitary Plan Wind Tunnels (UPWT) to determine such ascent aeroacoustic environments for the Space Launch System (SLS). Surface static pressure measurements were also collected to aid in determination of local environments for venting, CFD substantiation, and calibration of the flush air data system located on the launch abort system. Additionally, this test supported a NASA Engineering and Safety Center study of alternate booster nose caps. Testing occurred during two test campaigns: August - September 2013 and December 2013 - January 2014. Four primary model configurations were tested for ascent aeroacoustic environment definition. The SLS Block 1 vehicle was represented by a 2.5% full stack model and a 4% truncated model. Preliminary Block 1B payload and manned configurations were also tested, using 2.5% full stack and 4% truncated models respectively. This test utilized the 11 x 11 foot transonic and 9 x 7 foot supersonic tunnel sections at the ARC UPWT to collect data from Mach 0.7 through 2.5 at various total angles of attack. SLS Block 1 design environments were developed primarily using these data. SLS Block 1B preliminary environments have also been prepared using these data. This paper discusses the test and analysis methodology utilized, with a focus on the unsteady data collection and processing.

  15. Overview of the Novel Intelligent JAXA Active Rotor Program

    NASA Technical Reports Server (NTRS)

    Saito, Shigeru; Kobiki, Noboru; Tanabe, Yasutada; Johnson, Wayne; Yamauchi, Gloria K.; Young, Larry A.

    2010-01-01

    The Novel Intelligent JAXA Active Rotor (NINJA Rotor) program is a cooperative effort between JAXA and NASA, involving a test of a JAXA pressure-instrumented, active-flap rotor in the 40- by 80-Foot Wind Tunnel at Ames Research Center. The objectives of the program are to obtain an experimental database of a rotor with active flaps and blade pressure instrumentation, and to use that data to develop analyses to predict the aerodynamic and aeroacoustic performance of rotors with active flaps. An overview of the program is presented, including a description of the rotor and preliminary pretest calculations.

  16. Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines: Preprint

    SciTech Connect

    Migliore, P.; Oerlemans, S.

    2003-12-01

    Aeroacoustic tests of seven airfoils were performed in an open jet anechoic wind tunnel. Six of the airfoils are candidates for use on small wind turbines operating at low Reynolds number. One airfoil was tested for comparison to benchmark data. Tests were conducted with and without boundary layer tripping. In some cases a turbulence grid was placed upstream in the test section to investigate inflow turbulence noise. An array of 48 microphones was used to locate noise sources and separate airfoil noise from extraneous tunnel noise. Trailing edge noise was dominant for all airfoils in clean tunnel flow. With the boundary layer untripped, several airfoils exhibited pure tones that disappeared after proper tripping was applied. In the presence of inflow turbulence, leading edge noise was dominant for all airfoils.

  17. On vortex-airfoil interaction noise including span-end effects, with application to open-rotor aeroacoustics

    NASA Astrophysics Data System (ADS)

    Roger, Michel; Schram, Christophe; Moreau, Stéphane

    2014-01-01

    A linear analytical model is developed for the chopping of a cylindrical vortex by a flat-plate airfoil, with or without a span-end effect. The major interest is the contribution of the tip-vortex produced by an upstream rotating blade in the rotor-rotor interaction noise mechanism of counter-rotating open rotors. Therefore the interaction is primarily addressed in an annular strip of limited spanwise extent bounding the impinged blade segment, and the unwrapped strip is described in Cartesian coordinates. The study also addresses the interaction of a propeller wake with a downstream wing or empennage. Cylindrical vortices are considered, for which the velocity field is expanded in two-dimensional gusts in the reference frame of the airfoil. For each gust the response of the airfoil is derived, first ignoring the effect of the span end, assimilating the airfoil to a rigid flat plate, with or without sweep. The corresponding unsteady lift acts as a distribution of acoustic dipoles, and the radiated sound is obtained from a radiation integral over the actual extent of the airfoil. In the case of tip-vortex interaction noise in CRORs the acoustic signature is determined for vortex trajectories passing beyond, exactly at and below the tip radius of the impinged blade segment, in a reference frame attached to the segment. In a second step the same problem is readdressed accounting for the effect of span end on the aerodynamic response of a blade tip. This is achieved through a composite two-directional Schwarzschild's technique. The modifications of the distributed unsteady lift and of the radiated sound are discussed. The chained source and radiation models provide physical insight into the mechanism of vortex chopping by a blade tip in free field. They allow assessing the acoustic benefit of clipping the rear rotor in a counter-rotating open-rotor architecture.

  18. Spin test of turbine rotor

    NASA Technical Reports Server (NTRS)

    Vavra, M. H.; Hammer, J. E.; Bell, L. E.

    1972-01-01

    Experimental data are presented for the tangential and radial stresses in the disks of the 36,000 horsepower, 4000 rpm turbine for the M-1 engine oxidizer turbopump. The two-stage Curtis turbine is a special light-weight design utilizing thin conical disks with hollow sheet metal blades attached by electron-beam welding techniques. The turbine was fabricated from Inconel 718, a nickel-chromium alloy. The stresses were obtained by strain-gage measurements using a slip-ring assembly to transmit the electrical signals. Measurements were made at different rotative speeds and different thermal loads. In addition to presenting test data, the report describes test equipment, design of associated hardware, test procedures, instrumentation, and tests for the selection and calibration of strain gages.

  19. HARP model rotor test at the DNW. [Hughes Advanced Rotor Program

    NASA Technical Reports Server (NTRS)

    Dawson, Seth; Jordan, David; Smith, Charles; Ekins, James; Silverthorn, Lou

    1989-01-01

    Data from a test of a dynamically scaled model of the Hughes Advanced Rotor Program (HARP) bearingless model main rotor and 369K tail rotor are reported. The history of the HARP program and its goals are reviewed, and the main and tail rotor models are described. The test facilities and instrumentation are described, and wind tunnel test data are presented on hover, forward flight performance, and blade-vortex interaction. Performance data, acoustic data, and dynamic data from near field/far field and shear layer studies are presented.

  20. Full-Span Tiltrotor Aeroacoustic Model (TRAM) Overview and 40- by 80-Foot Wind Tunnel Test. [conducted in the 40- by 80-Foot Wind Tunnel at Ames Research Center

    NASA Technical Reports Server (NTRS)

    McCluer, Megan S.; Johnson, Jeffrey L.; Rutkowski, Michael (Technical Monitor)

    2001-01-01

    Most helicopter data trends cannot be extrapolated to tiltrotors because blade geometry and aerodynamic behavior, as well as rotor and fuselage interactions, are significantly different for tiltrotors. A tiltrotor model has been developed to investigate the aeromechanics of tiltrotors, to develop a comprehensive database for validating tiltrotor analyses, and to provide a research platform for supporting future tiltrotor designs. The Full-Span Tiltrotor Aeroacoustic Model (FS TRAM) is a dual-rotor, powered aircraft model with extensive instrumentation for measurement of structural and aerodynamic loads. This paper will present the Full-Span TRAM test capabilities and the first set of data obtained during a 40- by 80-Foot Wind Tunnel test conducted in late 2000 at NASA Ames Research Center. The Full-Span TRAM is a quarter-scale representation of the V-22 Osprey aircraft, and a heavily instrumented NASA and U.S. Army wind tunnel test stand. Rotor structural loads are monitored and recorded for safety-of-flight and for information on blade loads and dynamics. Left and right rotor balance and fuselage balance loads are monitored for safety-of-flight and for measurement of vehicle and rotor aerodynamic performance. Static pressure taps on the left wing are used to determine rotor/wing interactional effects and rotor blade dynamic pressures measure blade airloads. All of these measurement capabilities make the FS TRAM test stand a unique and valuable asset for validation of computational codes and to aid in future tiltrotor designs. The Full-Span TRAM was tested in the NASA Ames Research Center 40- by 80-Foot Wind Tunnel from October through December 2000. Rotor and vehicle performance measurements were acquired in addition to wing pressures, rotor acoustics, and Laser Light Sheet (LLS) flow visualization data. Hover, forward flight, and airframe (rotors off) aerodynamic runs were performed. Helicopter-mode data were acquired during angle of attack and thrust sweeps for

  1. Test results from a dynamic model dynaflex rotor

    NASA Technical Reports Server (NTRS)

    Niebanck, C. F.; Goodman, R. K.

    1985-01-01

    A one-fifth scale dynamic model of the Sikorsky Dynaflex rotor was tested in hover and in forward flight conditions in a wind tunnel. The Dynaflex rotor features an advanced composite structure which flexes to provide a constant speed universal joint action. Testing concentrated on confirming that the stability and dynamic response of the rotor were satisfactory. Lift conditions of up to .11 Ct/sigma and advance ratios as high as .46 were reached. Vibratory loads were compared to those of articulated rotors. The Dynaflex rotor concept appears to be a practical concept from the standpoint of dynamic response and stability.

  2. Icing research tunnel test of a model helicopter rotor

    NASA Technical Reports Server (NTRS)

    Miller, Thomas L.; Bond, Thomas H.

    1989-01-01

    An experimental program has been conducted in the NASA Lewis Research Center Icing Research Tunnel (IRT) in which an OH-58 tail rotor assembly was operated in a horizontal plane to simulate the action of a typical main rotor. Ice was accreted on the blades in a variety of rotor and tunnel operating conditions and documentation of the resulting shapes was performed. Rotor torque and vibration are presented as functions of time for several representative test runs, and the effects of various parametric variations on the blade ice shapes are shown. This OH-58 test was the first of its kind in the United States and will encourage additional model rotor icing tunnel testing. Although not a scaled representative of any actual full-scale main rotor system, this rig has produced torque and vibration data which will be useful in assessing the quality of existing rotor icing analyses.

  3. Psychoacoustic Testing of Modulated Blade Spacing for Main Rotors

    NASA Technical Reports Server (NTRS)

    Edwards, Bryan; Booth, Earl R., Jr. (Technical Monitor)

    2002-01-01

    Psychoacoustic testing of simulated helicopter main rotor noise is described, and the subjective results are presented. The objective of these tests was to evaluate the potential acoustic benefits of main rotors with modulated (uneven) blade spacing. Sound simulations were prepared for six main rotor configurations. A baseline 4-blade main rotor with regular blade spacing was based on the Bell Model 427 helicopter. A 5-blade main rotor with regular spacing was designed to approximate the performance of the 427, but at reduced tipspeed. Four modulated rotors - one with "optimum" spacing and three alternate configurations - were derived from the 5 bladed regular spacing rotor. The sounds were played to 2 subjects at a time, with care being taken in the speaker selection and placement to ensure that the sounds were identical for each subject. A total of 40 subjects participated. For each rotor configuration, the listeners were asked to evaluate the sounds in terms of noisiness. The test results indicate little to no "annoyance" benefit for the modulated blade spacing. In general, the subjects preferred the sound of the 5-blade regular spaced rotor over any of the modulated ones. A conclusion is that modulated blade spacing is not a promising design feature to reduce the annoyance for helicopter main rotors.

  4. V/STOL tilt rotor aircraft study: Wind tunnel tests of a full scale hingeless prop/rotor designed for the Boeing Model 222 tilt rotor aircraft

    NASA Technical Reports Server (NTRS)

    Magee, J. P.; Alexander, H. R.

    1973-01-01

    The rotor system designed for the Boeing Model 222 tilt rotor aircraft is a soft-in-plane hingeless rotor design, 26 feet in diameter. This rotor has completed two test programs in the NASA Ames 40' X 80' wind tunnel. The first test was a windmilling rotor test on two dynamic wing test stands. The rotor was tested up to an advance ratio equivalence of 400 knots. The second test used the NASA powered propeller test rig and data were obtained in hover, transition and low speed cruise flight. Test data were obtained in the areas of wing-rotor dynamics, rotor loads, stability and control, feedback controls, and performance to meet the test objectives. These data are presented.

  5. 14 CFR 33.92 - Rotor locking tests.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Rotor locking tests. 33.92 Section 33.92 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.92 Rotor locking tests. If...

  6. Flight Test Results for Uniquely Tailored Propulsion-Airframe Aeroacoustic Chevrons: Shockcell Noise

    NASA Technical Reports Server (NTRS)

    Mengle, Vinod G.; Ganz, Ulrich W.; Nesbitt, Eric; Bultemeier, Eric J.; Thomas, Russell H.; Nesbitt, Eric

    2006-01-01

    Azimuthally varying chevrons (AVC) which have been uniquely tailored to account for the asymmetric propulsion-airframe aeroacoustic interactions have recently shown significant reductions in jet-related community noise at low-speed take-off conditions in scale model tests of coaxial nozzles with high bypass ratio. There were indications that such AVCs may also provide shockcell noise reductions at high cruise speeds. This paper describes the flight test results when one such AVC concept, namely, the T-fan chevrons with enhanced mixing near the pylon, was tested at full-scale on a modern large twin-jet aircraft (777-300ER) with focus on shockcell noise at mid-cruise conditions. Shockcell noise is part of the interior cabin noise at cruise conditions and its reduction is useful from the viewpoint of passenger comfort. Noise reduction at the source, in the exhaust jet, especially, at low frequencies, is beneficial from the perspective of reduced fuselage sidewall acoustic lining. Results are shown in terms of unsteady pressure spectra both on the exterior surface of the fuselage at several axial stations and also microphone arrays placed inside the fuselage aft of the engine. The benefits of T-fan chevrons, with and without conventional chevrons on the core nozzle, are shown for several engine operating conditions at cruise involving supersonic fan stream and subsonic or supersonic core stream. The T-fan AVC alone provides up to 5 dB low-frequency noise reduction on the fuselage exterior skin and up to 2 dB reduction inside the cabin. Addition of core chevrons appears to increase the higher frequency noise. This flight test result with the previous model test observation that the T-fan AVCs have hardly any cruise thrust coefficient loss (< 0.05%) make them viable candidates for reducing interior cabin noise in high bypass ratio engines.

  7. Results of a sub-scale model rotor icing test

    NASA Technical Reports Server (NTRS)

    Flemming, Robert J.; Bond, Thomas H.; Britton, Randall K.

    1991-01-01

    A heavily instrumented sub-scale model of a helicopter main rotor was tested in the NASA Lewis Research Center Icing Research Tunnel (IRT) in September and November 1989. The four-bladed main rotor had a diameter of 1.83 m (6.00 ft) and the 0.124 m (4.9 in) chord rotor blades were specially fabricated for this experiment. The instrumented rotor was mounted on a Sikorsky Aircraft Powered Force Model, which enclosed a rotor balance and other measurement systems. The model rotor was exposed to a range of icing conditions that included variations in temperature, liquid water content, and median droplet diameter, and was operated over ranges of advance ratio, shaft angle, tip Mach number (rotor speed) and weight coefficient to determine the effect of these parameters on ice accretion. In addition to strain gage and balance data, the test was documented with still, video, and high speed photography, ice profile tracings, and ice molds. The sensitivity of the model rotor to the test parameters, is given, and the result to theoretical predictions are compared. Test data quality was excellent, and ice accretion prediction methods and rotor performance prediction methods (using published icing lift and drag relationships) reproduced the performance trends observed in the test. Adjustments to the correlation coefficients to improve the level of correlation are suggested.

  8. Wind-tunnel Tests of a Cyclogiro Rotor

    NASA Technical Reports Server (NTRS)

    Wheatley, John B; Windler, Ray

    1935-01-01

    During an extensive study of all types of rotating wings, the NACA examined the cyclogiro rotor and made an aerodynamic analysis of that system (reference 1). The examination disclosed that such a machine had sufficient promise to justify an experimental investigation; a model with a diameter and span of 8 feet was therefore constructed and tested in the 20-foot wind tunnel during 1934. The experimental work included tests of the effect of the motion upon the rotor forces during the static-lift and forward-flight conditions at several rotor speeds and the determination of the relations between the forces generated by the rotor and the power required by it.

  9. Flight Test Results for Uniquely Tailored Propulsion-Airframe Aeroacoustic Chevrons: Community Noise

    NASA Technical Reports Server (NTRS)

    Nesbitt, Eric; Mengle, Vinod; Czech, Michael; Callendar, Bryan; Thomas, Russ

    2006-01-01

    The flow/acoustic environment around the jet exhaust of an engine when installed on an airplane, say, under the wing, is highly asymmetric due to the pylon, the wing and the high-lift devices. Recent scale model tests have shown that such Propulsion Airframe Aeroacoustic (PAA) interactions and the jet mixing noise can be reduced more than with conventional azimuthally uniform chevrons by uniquely tailoring the chevrons to produce enhanced mixing near the pylon. This paper describes the community noise results from a flight test on a large twin-engine airplane using this concept of azimuthally varying chevrons for engines installed under the wing. Results for two different nozzle configurations are described: azimuthally varying "PAA T-fan" chevrons on the fan nozzle with a baseline no-chevron core nozzle and a second with PAA T-fan chevrons with conventional azimuthally uniform chevrons on the core nozzle. We analyze these test results in comparison to the baseline no-chevron nozzle on both spectral and integrated power level bases. The study focuses on the peak jet noise reduction and the effects at high frequencies for typical take-off power settings. The noise reduction and the absolute noise levels are then compared to model scale results. The flight test results verify that the PAA T-fan nozzles in combination with standard core chevron nozzles can, indeed, give a reasonable amount of noise reduction at low frequencies without high-frequency lift during take-off conditions and hardly any impact on the cruise thrust coefficient.

  10. Flow Measurements of an Isolated Model Tilt Rotor

    DTIC Science & Technology

    1999-05-01

    Murashige , A., Kobiki, N., Tsuchihashi, A., Nakamura, H., Inagaki, K., Yamakawa, E., “ATIC Aeroacoustic Model Rotor Test at DNW,” AHS International...Meeting on Advanced Rotorcraft Technology and Disaster Relief, Gifu, Japan, April 1998. 25. Murashige , A., Tsuchihashi, A., Tsujiuchi, T., Yamakawa, E

  11. Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines; Period of Performance: August 23, 2002 through March 31, 2004

    SciTech Connect

    Oerlemans, S.

    2004-08-01

    The U.S. Department of Energy, working through the National Renewable Energy Laboratory, is engaged in a comprehensive research effort to improve our understanding of wind turbine aeroacoustics. Quiet wind turbines are an inducement to widespread deployment, so the goal of NREL's aeroacoustic research is to develop tools that the U.S. wind industry can use in developing and deploying highly efficient, quiet wind turbines at low wind speed sites. NREL's National Wind Technology Center is implementing a multifaceted approach that includes wind tunnel tests, field tests, and theoretical analyses in direct support of low wind speed turbine development by its industry partners. To that end, wind tunnel aerodynamic tests and aeroacoustic tests have been performed on six airfoils that are candidates for use on small wind turbines. Results are documented in this report.

  12. SMART wind turbine rotor. Design and field test

    SciTech Connect

    Berg, Jonathan Charles; Resor, Brian Ray; Paquette, Joshua A.; White, Jonathan Randall

    2014-01-01

    The Wind Energy Technologies department at Sandia National Laboratories has developed and field tested a wind turbine rotor with integrated trailing-edge flaps designed for active control of rotor aerodynamics. The SMART Rotor project was funded by the Wind and Water Power Technologies Office of the U.S. Department of Energy (DOE) and was conducted to demonstrate active rotor control and evaluate simulation tools available for active control research. This report documents the design, fabrication, and testing of the SMART Rotor. This report begins with an overview of active control research at Sandia and the objectives of this project. The SMART blade, based on the DOE / SNL 9-meter CX-100 blade design, is then documented including all modifications necessary to integrate the trailing edge flaps, sensors incorporated into the system, and the fabrication processes that were utilized. Finally the test site and test campaign are described.

  13. SMART Rotor Development and Wind-Tunnel Test

    NASA Technical Reports Server (NTRS)

    Lau, Benton H.; Straub, Friedrich; Anand, V. R.; Birchette, Terry

    2009-01-01

    Boeing and a team from Air Force, NASA, Army, Massachusetts Institute of Technology, University of California at Los Angeles, and University of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center, figure 1. The SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing-edge flap on each blade. The development effort included design, fabrication, and component testing of the rotor blades, the trailing-edge flaps, the piezoelectric actuators, the switching power amplifiers, the actuator control system, and the data/power system. Development of the smart rotor culminated in a whirl-tower hover test which demonstrated the functionality, robustness, and required authority of the active flap system. The eleven-week wind tunnel test program evaluated the forward flight characteristics of the active-flap rotor, gathered data to validate state-of-the-art codes for rotor noise analysis, and quantified the effects of open- and closed-loop active-flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness and the reliability of the flap actuation system were successfully demonstrated in more than 60 hours of wind-tunnel testing. The data acquired and lessons learned will be instrumental in maturing this technology and transitioning it into production. The development effort, test hardware, wind-tunnel test program, and test results will be presented in the full paper.

  14. Hover performance tests of full scale variable geometry rotors

    NASA Technical Reports Server (NTRS)

    Rorke, J. B.

    1976-01-01

    Full scale whirl tests were conducted to determine the effects of interblade spatial relationships and pitch variations on the hover performance and acoustic signature of a 6-blade main rotor system. The variable geometry rotor (VGR) variations from the conventional baseline were accomplished by: (1) shifting the axial position of alternate blades by one chord-length to form two tip path planes; and (2) varying the relative azimuthal spacing from the upper rotor to the lagging hover rotor in four increments from 25.2 degrees to 62.1 degrees. For each of these four configurations, the differential collective pitch between upper and lower rotors was set at + or - 1 deg, 0 deg and -1 deg. Hover performance data for all configurations were acquired at blade tip Mach numbers of 0.523 and 0.45. Acoustic data were recorded at all test conditions, but analyzed only at 0 deg differential pitch at the higher rotor speed. The VGR configurations tested demonstrated improvements in thrust at constant power as high as 6 percent. Reductions of 3 PNdb in perceived noise level and of 4 db in blade passage frequency noise level were achieved at the higher thrust levels. Consistent correlation exists between performance and acoustic improvements. For any given azimuth spacing, performance was consistently better for the differential pitch condition of + or - 1 degree, i.e. with the upper rotor pitch one degree higher than the lower rotor.

  15. Design, analysis and testing of small, affordable HAWT rotors

    NASA Astrophysics Data System (ADS)

    Pricop, Mihai V.; Niculescu, Mihai L.; Cojocaru, Marius G.; Barsan, Dorin

    2012-09-01

    The paper presents affordable technologies dedicated to design, CAD modelling and manufacturing of the small-medium HAWT rotors. Three numerical tools are developed: blade/rotor design, blade modelling for industry CATIA(CATScript) and blade modelling for small scale developers. Numerical analysis of the rotors is accomplished for both performance and noise level estimation using XFLOW (LES) and an in-house code (URANS). Results are presented for a 5KW rotor at the design point only, since computations are expensive. Developement examples are included as two rotors are designed, manufactured and tested for 1.5 and 5KW. A third one, rated for 20KW is under developement. Basic testing results are also included.

  16. Rotor Performance at High Advance Ratio: Theory versus Test

    NASA Technical Reports Server (NTRS)

    Harris, Franklin D.

    2008-01-01

    Five analytical tools have been used to study rotor performance at high advance ratio. One is representative of autogyro rotor theory in 1934 and four are representative of helicopter rotor theory in 2008. The five theories are measured against three sets of well documented, full-scale, isolated rotor performance experiments. The major finding of this study is that the decades spent by many rotorcraft theoreticians to improve prediction of basic rotor aerodynamic performance has paid off. This payoff, illustrated by comparing the CAMRAD II comprehensive code and Wheatley & Bailey theory to H-34 test data, shows that rational rotor lift to drag ratios are now predictable. The 1934 theory predicted L/D ratios as high as 15. CAMRAD II predictions compared well with H-34 test data having L/D ratios more on the order of 7 to 9. However, the detailed examination of the selected codes compared to H-34 test data indicates that not one of the codes can predict to engineering accuracy above an advance ratio of 0.62 the control positions and shaft angle of attack required for a given lift. There is no full-scale rotor performance data available for advance ratios above 1.0 and extrapolation of currently available data to advance ratios on the order of 2.0 is unreasonable despite the needs of future rotorcraft. Therefore, it is recommended that an overly strong full-scale rotor blade set be obtained and tested in a suitable wind tunnel to at least an advance ratio of 2.5. A tail rotor from a Sikorsky CH-53 or other large single rotor helicopter should be adequate for this exploratory experiment.

  17. Hover performance tests of baseline metal and Advanced Technology Blade (ATB) rotor systems for the XV-15 tilt rotor aircraft

    NASA Technical Reports Server (NTRS)

    Bartie, K.; Alexander, H.; Mcveigh, M.; Lamon, S.; Bishop, H.

    1986-01-01

    Rotor hover performance data were obtained for two full-scale rotor systems designed for the XV-15 Tilt Rotor Research Aircraft. One rotor employed the rectangular planform metal blades (rotor solidity = 0.089) which were used on the initial flight configuration of the XV-15. The second rotor configuration examined the nonlinear taper, composite-construction, Advanced Technology Blade (ATB), (rotor solidity = 0.10) designed to replace the metal blades on the XV-15. Variations of the baseline ATB tip and cuff shapes were also tested. A new six-component rotor force and moment balance designed to obtain highly accurate data over a broad range of thrust and torque conditions is described. The test data are presented in nondimensional coefficient form for the performance results, and in dimensional form for the steady and alternating loads. Some wake and acoustic data are also shown.

  18. SMART Wind Turbine Rotor: Design and Field Test

    SciTech Connect

    Berg, Jonathan C.; Resor, Brian R.; Paquette, Joshua A.; White, Jonathan R.

    2014-01-29

    This report documents the design, fabrication, and testing of the SMART Rotor. This work established hypothetical approaches for integrating active aerodynamic devices (AADs) into the wind turbine structure and controllers.

  19. Acoustic test of a model rotor and tail rotor: Results for the isolated rotors and combined configuration

    NASA Technical Reports Server (NTRS)

    Martin, R. M.; Burley, C. L.; Elliott, J. W.

    1989-01-01

    Acoustic data from a model scale main rotor and tail rotor experiment in the NASA Langley 14 by 22 Foot Subsonic Tunnel are presented for the main rotor and trail rotor in isolation and for the two rotors operating together. Results for the isolated main rotor show the importance of the rotor flapping conditions on mid-frequency noise content. High levels of main rotor retreating side blade-vortex interaction noise are shown to radiate downstream of the model. The isolated tail rotor noise results show the dominance of harmonic noise in the thrusting direction. The occurrence of tail rotor broadband noise is seen by the broadening of the tail rotor harmonics and is attributed to fuselage wake turbulence. The combined main and tail rotor data are presented to show the dominance of each rotor's different noise sources at different directivity locations.

  20. Aeroacoustics - Historical perspective and important issues

    NASA Astrophysics Data System (ADS)

    Leverton, John W.

    Theoretical and experimental studies of helicopter rotor noise are reviewed, covering the period 1962-1987. Topics examined include the increased importance of rotor-induced noise with the advent of gas-turbine-powered helicopters in the early 1960s, the impact of improved acoustic instruments, harmonic analyses of rotor noise, aeroacoustic theory, rotational noise, fluctuating forces, broadband noise, blade slap, blade-vortex interaction, thickness noise, tail-rotor noise, and main-rotor-wake/tail-rotor interaction. It is suggested that the tip speed and number of blades of rotors for large helicopters can be selected to minimize noise at little loss of performance, but that this is not the case for smaller helicopters. In the latter case, greater attention to operational aspects in the overall design and greater cooperation between aerodynamicists and acoustic engineers are recommended.

  1. Aeroacoustic wind-tunnel tests on full-scale propellers with unsymmetrical blade spacing

    NASA Astrophysics Data System (ADS)

    Dobrzynski, W.

    In order to confirm predicted noise reductions for aircraft propellers designed with irregular blade spacing an aeroacoustic experiment is conducted on a 6-blade propeller. The two spacing angles between the pairs of blades are varied but kept identical, and noise spectra are measured along with polar noise-level directivities with comparisons to theoretical results. A maximum noise reduction of about 3 dB is found for helical-blade-tip Mach numbers in excess of 0.7 and blade-spacing angles in the range of 15-20 deg.

  2. Analysis of SRB reentry acoustic environments. [aeroacoustic spectra determined from wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Coffin, T.; Dandridge, R. E.; Haddock, U. W.

    1979-01-01

    Space shuttle solid rocket booster reentry aeroacoustic environments were estimated. Particular emphasis was given to the aft skirt/exit plane region for the Mach number regime 0.6 = or greater than M infinity = or less than 3.5. The analysis is based on the evaluation of wind tunnel model results in conjunction with Monte Carlo simulation of trajectory parameters. The experimental approach is described as well as the evaluation process utilized. Predicted environments are presented in terms of one-third octave band spectra representing space averaged values for critical regions on the solid rocket booster.

  3. Report on the final panel discussion on computational aeroacoustics

    NASA Technical Reports Server (NTRS)

    Lighthill, James

    1992-01-01

    Some important conclusions about future prospects for aeroacoustics in general, and for computational aeroacoustics in particular, that were reached in the course of the Final Panel Discussion of the Workshop on Computational Aeroacoustics held from 6 to 9 April 1992 by ICASE and NASA Langley Research Center are summarized by the panel chairman. Aeroacoustics must now be involved in interactions with computational fluid dynamics (as applied not only to deterministic flows but also to the statistical characteristics of turbulence), while additionally incorporating rigorous comparisons with experiment. The new Computational Aeroacoustics will press forward in two parallel ways. In one of them, CFD will be used to determine aeroacoustic source strengths, the associated radiation being derived by the Acoustic Analogy approach in one of its forms. In the other, a direct Computational Aeroacoustics will apply CFD techniques over a region extending beyond the flow field so as to include at least the beginnings of the acoustic far field. There are some particularly important areas of study, including rotor noise, boundary-layer noise, and the noise of supersonic jets, where it is strongly recommended that use of both methods is continued. On the other hand, important problems of the diffraction of radiation from aeroacoustic sources around complicated aircraft shapes will require the use of comprehensively Computational Aeroacoustics, while Acoustic Analogy methods seem better suited to estimating subsonic jet noise. The study of model problems to allow comparisons with experiment will be valuable in both lines of attack.

  4. Growian rotor blades: Production development, construction and test

    NASA Technical Reports Server (NTRS)

    Thiele, H. M.

    1984-01-01

    Development and construction of three 50 m rotor blades for a 3 MW wind turbine are described. A hybrid concept was chosen, i.e., a load carrying inflexible steel spar and a glass fiber reinforced plastic skin. A test blade was constructed and static loading tests, dynamic vibration tests and fatigue tests on critical welds as well as at the connection between spar and blade skin were performed. All test results show good accordance with calculated values, and were taken into consideration during the construction of two rotor blades.

  5. Aeroacoustic Duster

    NASA Technical Reports Server (NTRS)

    Marshall, Jeffrey S. (Inventor); Chen, Di (Inventor); Vachon, Nicholas Mario (Inventor); Hitt, Darren (Inventor); Wu, Junru (Inventor)

    2014-01-01

    The aero-acoustic duster invention disclosed herein provides for high particle removal rate from surfaces with low energy expenditure relative to competing vacuum-based devices. The device removes particulate matter from a surface using a two-step process: 1. Acoustic radiation is used to break the adhesive bonds between dust and the surface, forcing particles into a mode where they continuously bounce up and down on the surface; and, 2. A bounded vortex is generated over the surface, with suction in the vortex center and jets for blowing air along the periphery. The jets are tilted in the tangential direction to induce vortex motion within the suction region. The vortex is said to be bounded because streamlines originating in the downward jets are entrained back into the central vortex.

  6. 14 CFR 29.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Rotor drive system and control mechanism... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Rotor Drive System § 29.923 Rotor drive system and control mechanism tests. (a) Endurance tests, general. Each rotor...

  7. 14 CFR 29.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Rotor drive system and control mechanism... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Rotor Drive System § 29.923 Rotor drive system and control mechanism tests. (a) Endurance tests, general. Each rotor...

  8. Wind tunnel test on a 1/4.622 Froude scale, hingeless rotor, tilt rotor model, volume 1

    NASA Technical Reports Server (NTRS)

    Magee, J. P.; Alexander, H. R.

    1976-01-01

    Wing tunnel test data on a 1/4.622 Froude scale, hingeless rotor, tilt rotor mode are reported for all potential flight conditions through hover and a wide envelope of transitions. A mathematical model was used to describe the rotor system in real time simulation by means of regression analyses. Details of the model, test program and data system are provided together with four data files for hover and transition.

  9. Application of Three-Component PIV to a Hovering Rotor Wake

    DTIC Science & Technology

    2000-05-01

    Murashige , A., Tsuchihashi, A., Tsujiuchi, T., Yamakawa, E., “Blade-Tip Vortex Measurement by PIV,” 23rd European Rotorcraft Forum, Dresden, Germany...1998. 10. Murashige , A., Kobiki, N., Tsuchihashi, A., Nakamura, H., Inagaki, K., Yamakawa, E., “ATIC Aeroacoustic Model Rotor Test at DNW,” AHS

  10. The design of test-section inserts for higher speed aeroacoustic testing in the Ames 80- by 120-Foot Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Soderman, Paul T.; Olson, Larry E.

    1992-01-01

    An engineering feasibility study was made of aeroacoustic inserts designed for large-scale acoustic research on aircraft models in the 80- by 120 Foot Wind Tunnel at NASA Ames Research Center. The goal was to find test-section modifications that would allow improved aeroacoustic testing at airspeeds equal to and above the current 100 knots limit. Results indicate that the required maximum airspeed drives the design of a particular insert. Using goals of 200, 150, and 100 knots airspeed, the analysis led to a 30 x 60 ft open-jet test section, a 40 x 80 ft open-jet test section, and a 70 x 110 ft closed test section with enhanced wall lining respectively. The open-jet inserts would be composed of a nozzle, collector, diffuser, and acoustic wedges incorporated in the existing 80 x 120 ft test section. The closed test section would be composed of approximately 5-ft acoustic wedges covered by a porous plate attached to the test-section walls of the existing 80 x 120. All designs would require a double row of acoustic vanes between the test section and fan drive to attenuate fan noise and, in the case of the open-jet designs, to control flow separation at the diffuser downstream end. The inserts would allow virtually anechoic acoustics studies of large helicopter models, jets and V/STOL aircraft models in simulated flight. Model scale studies would be necessary to optimize the aerodynamic and acoustic performance of any of the designs.

  11. Fan Noise Source Diagnostic Test: Rotor Alone Aerodynamic Performance Results

    NASA Technical Reports Server (NTRS)

    Hughes, Christopher E.; Jeracki, Robert J.; Woodward, Richard P.; Miller, Christopher J.

    2005-01-01

    The aerodynamic performance of an isolated fan or rotor alone model was measured in the NASA Glenn Research Center 9- by 15- Foot Low Speed Wind Tunnel as part of the Fan Broadband Source Diagnostic Test conducted at NASA Glenn. The Source Diagnostic Test was conducted to identify the noise sources within a wind tunnel scale model of a turbofan engine and quantify their contribution to the overall system noise level. The fan was part of a 1/5th scale model representation of the bypass stage of a current technology turbofan engine. For the rotor alone testing, the fan and nacelle, including the inlet, external cowl, and fixed area fan exit nozzle, were modeled in the test hardware; the internal outlet guide vanes located behind the fan were removed. Without the outlet guide vanes, the velocity at the nozzle exit changes significantly, thereby affecting the fan performance. As part of the investigation, variations in the fan nozzle area were tested in order to match as closely as possible the rotor alone performance with the fan performance obtained with the outlet guide vanes installed. The fan operating performance was determined using fixed pressure/temperature combination rakes and the corrected weight flow. The performance results indicate that a suitable nozzle exit was achieved to be able to closely match the rotor alone and fan/outlet guide vane configuration performance on the sea level operating line. A small shift in the slope of the sea level operating line was measured, which resulted in a slightly higher rotor alone fan pressure ratio at take-off conditions, matched fan performance at cutback conditions, and a slightly lower rotor alone fan pressure ratio at approach conditions. However, the small differences in fan performance at all fan conditions were considered too small to affect the fan acoustic performance.

  12. NASA/GE Collaboration on Open Rotors - High Speed Testing

    NASA Technical Reports Server (NTRS)

    VanZante, Dale E.

    2011-01-01

    A low-noise open rotor system is being tested in collaboration with General Electric and CFM International, a 50/50 joint company between Snecmaand GE. Candidate technologies for lower noise will be investigated as well as installation effects such as pylon integration. Current test status for the 8x6 SWT high speed testing is presented as well as future scheduled testing which includes the FAA/CLEEN test entry. The tunnel blockage and propeller thrust calibration configurations are shown.

  13. Reducing the Effect of Transducer Mount Induced Noise on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

    NASA Technical Reports Server (NTRS)

    Herron, A. J.; Reed, D. K.; Nance, D. K.

    2015-01-01

    Characterization of launch vehicle unsteady aerodynamics is a field best studied through experimentation, which is often carried out in the form of large scale wind tunnel testing. Measurement of the fluctuating pressures induced by the boundary layer noise is customarily made with miniature pressure transducers installed into a model of the vehicle of interest. Literature shows that noise level increases between two to five decibels (dB referenced to 20 micropascal) can be induced when the transducer surface is not mounted perfectly flush with the model outer surface. To reduce this artificially induced noise, special transducer holders have been used for aeroacoustic wind tunnel testing by NASA. This holder is a sleeve into which the transducer fits, with a cap that allows it to be mounted in a recessed hole in the model. A single hole in the holder allows the transport of the tunnel medium so the transducer can discriminate the fluctuating pressure due to the turbulent boundary layer noise. The holder is first dry fitted into the model and any difference in height between the holder and the model surface can be sanded flush. The holder is then removed from the model, the transducer glued inside the holder, and the holder replaced in the model, secured also with glue, thus eliminating the problem of noise level increases due to lack of flushness. In order to work with this holder design, special transducers have been ordered with their standard screen removed and the diaphragm moved as close to the top of the casing as possible to minimize any cavity volume. Although this greatly reduces induced noise due to the transducers being out of flush, the holders can also induce a cavity resonance that is usually at a very high frequency. This noise is termed transducer mount induced noise (XMIN). The peak of the mode can vary with the cavity depth, boundary layer noise that can excite the mode, tunnel flow medium, and the build of the transducers. Because the boundary

  14. Hover test of a full-scale hingeless helicopter rotor: Aeroelastic stability, performance and loads data. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Peterson, R. L.; Warmbrodt, W.

    1984-01-01

    A hover test of a full-scale, hingeless rotor system was conducted in the NASA Ames 40- by 80-foot wind tunnel. The rotor was tested on the Ames rotor test apparatus. Rotor aeroelastic stability, performance, and loads at various rotational speeds and thrust coefficients were investigated. The primary objective was to determine the inplane stability characteristics of the rotor system. Rotor inplane damping data were obtained for operation between 350 and 425 rpm (design speed), and for thurst coefficients between 0.0 and 0.12. The rotor was stable for all conditions tested. At constant rotor rotational speed, a minimum inplane dampling level was obtained at a thrust coefficient approximately = 0.02. At constant rotor lift, a minimum in rotor inplane damping was measured at 400 rpm.

  15. Aeroacoustics of advanced propellers

    NASA Technical Reports Server (NTRS)

    Groeneweg, John F.

    1990-01-01

    The aeroacoustics of advanced, high speed propellers (propfans) are reviewed from the perspective of NASA research conducted in support of the Advanced Turboprop Program. Aerodynamic and acoustic components of prediction methods for near and far field noise are summarized for both single and counterrotation propellers in uninstalled and configurations. Experimental results from tests at both takeoff/approach and cruise conditions are reviewed with emphasis on: (1) single and counterrotation model tests in the NASA Lewis 9 by 15 (low speed) and 8 by 6 (high speed) wind tunnels, and (2) full scale flight tests of a 9 ft (2.74 m) diameter single rotation wing mounted tractor and a 11.7 ft (3.57 m) diameter counterrotation aft mounted pusher propeller. Comparisons of model data projected to flight with full scale flight data show good agreement validating the scale model wind tunnel approach. Likewise, comparisons of measured and predicted noise level show excellent agreement for both single and counterrotation propellers. Progress in describing angle of attack and installation effects is also summarized. Finally, the aeroacoustic issues associated with ducted propellers (very high bypass fans) are discussed.

  16. Reducing the Effect of Transducer Mount Induced Noise on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

    NASA Technical Reports Server (NTRS)

    Herron, Andrew J.; Reed, Darren K.; Nance, Donald K.

    2015-01-01

    Flight vehicle aeroacoustic environments induced during transonic and supersonic flight are usually predicted by subscale wind tunnel testing utilizing high frequency miniature pressure transducers. In order to minimize noise induced by the measurement itself, transducer flush mounting with the model surface is very important. The National Aeronautics and Space Administration (NASA) has accomplished flushness in recent testing campaigns via use of a transducer holder that can be machined and sanded. A single hole in the holder allows the flow medium to interact with the transducer diaphragm. Noise is induced by the resulting cavity however, and is a challenge to remove in post-processing. A new holder design has been developed that minimizes the effects of this transducer mount induced noise (XMIN) by reducing the resonance amplitude or increasing its resonance frequency beyond the range of interest. This paper describes a test conducted at the NASA/George C. Marshall Space Flight Center Trisonic Wind Tunnel intended to verify the effectiveness of this design. The results from this test show that this new transducer holder design does significantly reduce the influence of XMIN on measured fluctuating pressure levels without degrading a transducer's ability to accurately measure the noise external to the model.

  17. Shake Test Results and Dynamic Calibration Efforts for the Large Rotor Test Apparatus

    NASA Technical Reports Server (NTRS)

    Russell, Carl R.

    2014-01-01

    A shake test of the Large Rotor Test Apparatus (LRTA) was performed in an effort to enhance NASAscapability to measure dynamic hub loads for full-scale rotor tests. This paper documents the results of theshake test as well as efforts to calibrate the LRTA balance system to measure dynamic loads.Dynamic rotor loads are the primary source of vibration in helicopters and other rotorcraft, leading topassenger discomfort and damage due to fatigue of aircraft components. There are novel methods beingdeveloped to reduce rotor vibrations, but measuring the actual vibration reductions on full-scale rotorsremains a challenge. In order to measure rotor forces on the LRTA, a balance system in the non-rotatingframe is used. The forces at the balance can then be translated to the hub reference frame to measure therotor loads. Because the LRTA has its own dynamic response, the balance system must be calibrated toinclude the natural frequencies of the test rig.

  18. Acoustic measurements from a rotor blade-vortex interaction noise experiment in the German-Dutch Wind Tunnel (DNW)

    NASA Technical Reports Server (NTRS)

    Martin, Ruth M.; Splettstoesser, W. R.; Elliott, J. W.; Schultz, K.-J.

    1988-01-01

    Acoustic data are presented from a 40 percent scale model of the 4-bladed BO-105 helicopter main rotor, measured in the large European aeroacoustic wind tunnel, the DNW. Rotor blade-vortex interaction (BVI) noise data in the low speed flight range were acquired using a traversing in-flow microphone array. The experimental apparatus, testing procedures, calibration results, and experimental objectives are fully described. A large representative set of averaged acoustic signals is presented.

  19. The design of test-section inserts for higher speed aeroacoustic testing in the Ames 80- by 120-foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Soderman, Paul T.; Olson, Larry E.

    1992-01-01

    An engineering feasibility study was made of aeroacoustic inserts designed for large-scale acoustic research on aircraft models in the 80- by 120-Foot Wind Tunnel at NASA Ames Research Center. The goal was to find test-section modifications that would allow improved aeroacoustic testing at airspeeds equal to and above the current 100 knots limit. Results indicate that the required maximum airspeed drives the design of a particular insert. Using goals of 200, 150, and 100 knots airspeed, the analysis led to a 30 x 60 ft open-jet test section, a 40 x 80 ft open-jet test section, and a 70 x 110 ft closed test section with enhanced wall lining respectively. The open-jet inserts would be composed of a nozzle, collector, diffuser, and acoustic wedges incorporated in the existing 80 x 120 ft test section. The closed test section would be composed of approximately 5-ft acoustic wedges covered by a porous plate attached to the test-section walls of the existing 80 x 120. All designs would require a double row of acoustic vanes between the test section and fan drive to attenuate fan noise and, in the case of the open-jet designs, to control flow separation at the diffuser downstream end. The inserts would allow virtually anechoic acoustics studies of large helicopter models, jets, and V/STOL aircraft models in simulated flight. Model scale studies would be necessary to optimize the aerodynamic and acoustic performance of any of the designs. Successful development of acoustically transparent walls, though not strictly necessary to the project, would lead to a porous-wall test section that could be substituted for any of the open-jet designs, and thereby eliminate many aerodynamic and acoustic problems characteristic of open-jet shear layers.

  20. Investigation of rotor blade element airloads for a teetering rotor in the blade stall regime (second wind tunnel test)

    NASA Technical Reports Server (NTRS)

    Dadone, L. U.; Fukushima, T.

    1975-01-01

    A test was conducted in the NASA-Ames 7 x 10 ft low speed wind tunnel on a seven-foot diameter model of a teetering rotor. The objectives of the test were: (1) acquire pressure data for correlation with laser and flow visualization measurements; (2) explore rotor propulsive force limits by varying the advance ratio at constant lift and propulsive force coefficients; (3) obtain additional data to define the differences between teetering and articulated rotors; and (4) verify the acceleration sensitivity of experimental transducers. Results are presented.

  1. A Static Burst Test for Composite Flywheel Rotors

    NASA Astrophysics Data System (ADS)

    Hartl, Stefan; Schulz, Alexander; Sima, Harald; Koch, Thomas; Kaltenbacher, Manfred

    2016-06-01

    High efficient and safe flywheels are an interesting technology for decentralized energy storage. To ensure all safety aspects, a static test method for a controlled initiation of a burst event for composite flywheel rotors is presented with nearly the same stress distribution as in the dynamic case, rotating with maximum speed. In addition to failure prediction using different maximum stress criteria and a safety factor, a set of tensile and compressive tests is carried out to identify the parameters of the used carbon fiber reinforced plastics (CFRP) material. The static finite element (FE) simulation results of the flywheel static burst test (FSBT) compare well to the quasistatic FE-simulation results of the flywheel rotor using inertia loads. Furthermore, it is demonstrated that the presented method is a very good controllable and observable possibility to test a high speed flywheel energy storage system (FESS) rotor in a static way. Thereby, a much more expensive and dangerous dynamic spin up test with possible uncertainties can be substituted.

  2. Laser doppler velocimeter system for subsonic jet mixer nozzle testing at the NASA Lewis Aeroacoustic Propulsion Lab

    NASA Technical Reports Server (NTRS)

    Podboy, Gary G.; Bridges, James E.; Saiyed, Naseem H.; Krupar, Martin J.

    1995-01-01

    A laser Doppler velocimeter (LDV) system developed for the Aeroacoustic Propulsion Laboratory (APL) at the NASA Lewis Research Center is described. This system was developed to acquire detailed flow field data which could be used to quantify the effectiveness of internal exhaust gas mixers (IEGM's) and to verify and calibrate computational codes. The LDV was used as an orthogonal, three component system to measure the flow field downstream of the exit of a series of IEGM's and a reference axisymmetric splitter configuration. The LDV system was also used as a one component system to measure the internal axial flow within the nozzle tailpipe downstream of the mixers. These IEGM's were designed for low-bypass ratio turbofan engines. The data were obtained at a simulated low flight speed, high-power operating condition. The optical, seeding, and data acquisition systems of the LDV are described in detail. Sample flow field measurements are provided to illustrate the capabilities of the system at the time of this test, which represented the first use of LDV at the APL. A discussion of planned improvements to the LDV is also included.

  3. 14 CFR 27.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Rotor drive system and control mechanism... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Rotor Drive System § 27.923 Rotor drive system and control mechanism tests. (a) Each part tested as prescribed in this...

  4. 14 CFR 27.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Rotor drive system and control mechanism... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Rotor Drive System § 27.923 Rotor drive system and control mechanism tests. (a) Each part tested as prescribed in this...

  5. 14 CFR 27.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Rotor drive system and control mechanism... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Rotor Drive System § 27.923 Rotor drive system and control mechanism tests. (a) Each part tested as prescribed in this...

  6. 14 CFR 27.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Rotor drive system and control mechanism... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Rotor Drive System § 27.923 Rotor drive system and control mechanism tests. (a) Each part tested as prescribed in this...

  7. 14 CFR 27.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Rotor drive system and control mechanism... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Rotor Drive System § 27.923 Rotor drive system and control mechanism tests. (a) Each part tested as prescribed in this...

  8. Analysis and correlation with theory of rotor lift-limit test data

    NASA Technical Reports Server (NTRS)

    Sheffler, M.

    1979-01-01

    A wind tunnel test program to define the cruise performance and determine any limitations to lift and propulsive force of a conventional helicopter rotor is described. A 2.96 foot radius model rotor was used. The maximum lift and propulsive force obtainable from an articulated rotor for advance ratios of 0.4 to 0.67, and the blade load growth as the lift approaches the limit are determined. Cruise rotor performance for advance ratios of 0.4 to 0.67 and the sensitivity of the rotor forces and moments to rotor control inputs as the lift limit is approached are established.

  9. Parametric Blade Study Test Report Rotor Configuration. Number 2

    DTIC Science & Technology

    1988-11-01

    by block number) FIELD GROUP SUB. GR. Axial compressor 01 -Gas Turbine t Aircraft Turbine Engine . 19, BSTRACT (Continue on reverse if nee..ary and...AFWAL-TR-88-41108 00 PARAMETRIC BLADE STUDY TEST REPORT 00 ROTOR CONFIGURATION NO. 2 N C. Herbert Law Steven L. Puterbaugh Technology Branch Turbine ...Technology Branch Technology Branch Turbine Engine Division FOR THE COMMANDER THOMAS J. IMS, Director Turbinengine Division Aero P opulsion Laboratory If

  10. Parametric Blade Study Test Report Rotor Configuration. Number 1

    DTIC Science & Technology

    1988-11-01

    Lauderdale, Florida) provides a visual indication of rotor speed accurate to ten RPM. The tachometer also includes an adjustable speed limiting switch as...elevated power requirements for the drive motor at 90-100% of design speed caused the motor windings to overheat. To permit adequate test time at these...Puterbaugh Technology Branch Turbine Engine Division November 1988 Interim Report for Period 1 January 1987 - 31 May 1988 Distribution to U.S. government

  11. Fabrication and testing of prestressed composite rotor blade spar specimens

    NASA Technical Reports Server (NTRS)

    Gleich, D.

    1974-01-01

    Prestressed composite spar specimens were fabricated and evaluated by crack propagation and ballistic penetration tests. The crack propagation tests on flawed specimens showed that the prestressed composite spar construction significantly suppresses crack growth. Damage from three high velocity 30 caliber projectile hits was confined to three small holes in the ballistic test specimen. No fragmentation or crack propagation was observed indicating good ballistic damage resistance. Rotor attachment approaches and improved structural performance configurations were identified. Design theory was verified by tests. The prestressed composite spar configuration consisted of a compressively prestressed high strength ARDEFORM 301 stainless steel liner overwrapped with pretensioned S-994 fiberglass.

  12. Aeroacoustics of Space Vehicles

    NASA Technical Reports Server (NTRS)

    Panda, Jayanta

    2014-01-01

    While for airplanes the subject of aeroacoustics is associated with community noise, for space vehicles it is associated with vibro-acoustics and structural dynamics. Surface pressure fluctuations encountered during launch and travel through lower part of the atmosphere create intense vibro-acoustics environment for the payload, electronics, navigational equipment, and a large number of subsystems. All of these components have to be designed and tested for flight-certification. This presentation will cover all three major sources encountered in manned and unmanned space vehicles: launch acoustics, ascent acoustics and abort acoustics. Launch pads employ elaborate acoustic suppression systems to mitigate the ignition pressure waves and rocket plume generated noise during the early part of the liftoff. Recently we have used large microphone arrays to identify the noise sources during liftoff and found that the standard model by Eldred and Jones (NASA SP-8072) to be grossly inadequate. As the vehicle speeds up and reaches transonic speed in relatively denser part of the atmosphere, various shock waves and flow separation events create unsteady pressure fluctuations that can lead to high vibration environment, and occasional coupling with the structural modes, which may lead to buffet. Examples of wind tunnel tests and computational simulations to optimize the outer mold line to quantify and reduce the surface pressure fluctuations will be presented. Finally, a manned space vehicle needs to be designed for crew safety during malfunctioning of the primary rocket vehicle. This brings the subject of acoustic environment during abort. For NASAs Multi-Purpose Crew Vehicle (MPCV), abort will be performed by lighting rocket motors atop the crew module. The severe aeroacoustics environments during various abort scenarios were measured for the first time by using hot helium to simulate rocket plumes in the Ames unitary plan wind tunnels. Various considerations used for the

  13. Wind-Tunnel Tests of a 10-foot-diameter Gyroplane Rotor

    NASA Technical Reports Server (NTRS)

    Wheatley, John B; Bioletti, Carlton

    1936-01-01

    This report presents the results of wind-tunnel tests on a model gyroplane rotor 10 feet in diameter. The rotor blades had zero sweepback and zero offset; the hub contained a feathering mechanism that provided control of the rotor rolling moment, but not of the pitching moment. The rotor was tested with 4 blades and with 2 blades. The entire useful range of pitch settings and tip-speed ratios was investigated including the phase of operation in which the rotor turned very slowly, or idled.

  14. Important Scaling Parameters for Testing Model-Scale Helicopter Rotors

    NASA Technical Reports Server (NTRS)

    Singleton, Jeffrey D.; Yeager, William T., Jr.

    1998-01-01

    An investigation into the effects of aerodynamic and aeroelastic scaling parameters on model scale helicopter rotors has been conducted in the NASA Langley Transonic Dynamics Tunnel. The effect of varying Reynolds number, blade Lock number, and structural elasticity on rotor performance has been studied and the performance results are discussed herein for two different rotor blade sets at two rotor advance ratios. One set of rotor blades were rigid and the other set of blades were dynamically scaled to be representative of a main rotor design for a utility class helicopter. The investigation was con-densities permits the acquisition of data for several Reynolds and Lock number combinations.

  15. Open Rotor Noise Shielding by Blended-Wing-Body Aircraft

    NASA Technical Reports Server (NTRS)

    Guo, Yueping; Czech, Michael J.; Thomas, Russell H.

    2015-01-01

    This paper presents an analysis of open rotor noise shielding by Blended Wing Body (BWB) aircraft by using model scale test data acquired in the Boeing Low Speed Aeroacoustic Facility (LSAF) with a legacy F7/A7 rotor model and a simplified BWB platform. The objective of the analysis is the understanding of the shielding features of the BWB and the method of application of the shielding data for noise studies of BWB aircraft with open rotor propulsion. By studying the directivity patterns of individual tones, it is shown that though the tonal energy distribution and the spectral content of the wind tunnel test model, and thus its total noise, may differ from those of more advanced rotor designs, the individual tones follow directivity patterns that characterize far field radiations of modern open rotors, ensuring the validity of the use of this shielding data. Thus, open rotor tonal noise shielding should be categorized into front rotor tones, aft rotor tones and interaction tones, not only because of the different directivities of the three groups of tones, but also due to the differences in their source locations and coherence features, which make the respective shielding characteristics of the three groups of tones distinctly different from each other. To reveal the parametric trends of the BWB shielding effects, results are presented with variations in frequency, far field emission angle, rotor operational condition, engine installation geometry, and local airframe features. These results prepare the way for the development of parametric models for the shielding effects in prediction tools.

  16. Performance and loads data from an outdoor hover test of a Lynx tail rotor

    NASA Technical Reports Server (NTRS)

    Signor, David B.; Yamauchi, Gloria K.; Smith, Charles A.; Hagen, Martin J.

    1989-01-01

    A Lynx tail rotor was tested in hover at the Outdoor Aerodynamic Research Facility at NASA Ames Research Center. The test objectives were to measure the isolated rotor performance to provide a baseline for subsequent testing, and to operate the rotor throughout the speed and collective envelope before testing in the NFAC 40- by 80-Foot Wind Tunnel. Rotor forces and blade bending moments were measured at ambient wind conditions from zero to 6.23 m/sec. The test envelope was limited to rotor speeds of 1550 to 1850 rpm and minus 13 deg to plus 20 deg of blade collective pitch. The isolated rotor performance and blade loads data are presented.

  17. Tests of Full-Scale Helicopter Rotors at High Advancing Tip Mach Numbers and Advance Ratios

    NASA Technical Reports Server (NTRS)

    Biggers, James C.; McCloud, John L., III; Stroub, Robert H.

    2015-01-01

    As a continuation of the studies of reference 1, three full-scale helicopter rotors have been tested in the Ames Research Center 40- by SO-foot wind tunnel. All three of them were two-bladed, teetering rotors. One of the rotors incorporated the NACA 0012 airfoil section over the entire length of the blade. This rotor was tested at advance ratios up to 1.05. Both of the other rotors were tapered in thickness and incorporated leading-edge camber over the outer 20 percent of the blade radius. The larger of these rotors was tested at advancing tip Mach numbers up to 1.02. Data were obtained for a wide range of lift and propulsive force, and are presented without discussion.

  18. Advanced radial inflow turbine rotor program: Design and dynamic testing

    NASA Technical Reports Server (NTRS)

    Rodgers, C.

    1976-01-01

    The advancement of small, cooled, radial inflow turbine technology in the area of operation at higher turbine inlet temperature is discussed. The first step was accomplished by designing, fabricating, and subjecting to limited mechanical testing an advanced gas generator rotating assembly comprising a radial inflow turbine and two-stage centrifugal compressor. The radial inflow turbine and second-stage compressor were designed as an integrally machined monorotor with turbine cooling taking place basically by conduction to the compressor. Design turbine inlet rotor gas temperature, rotational speed, and overall gas generator compressor pressure ratio were 1422 K (2560 R), 71,222 rpm, and 10/1 respectively. Mechanical testing on a fabricated rotating assembly and bearing system covered 1,000 cold start/stop cycles and three spins to 120 percent design speed (85,466 rpm).

  19. Performance and loads data from a hover test of a 0.658-scale V-22 rotor and wing

    NASA Technical Reports Server (NTRS)

    Felker, Fort F.; Signor, David B.; Young, Larry A.; Betzina, Mark D.

    1987-01-01

    A hover test of a 0.658-scale model of a V-22 rotor and wing was conducted at the Outdoor Aerodynamic Research Facility at Ames Research Center. The primary objectives of the test were to obtain accurate measurements of the hover performance of the rotor system, and to measure the aerodynamic interactions between the rotor and wing. Data were acquired for rotor tip Mach numbers ranging from 0.1 to 0.73. This report presents data on rotor performance, rotor-wake downwash velocities, rotor system loads, wing forces and moments, and wing surface pressures.

  20. Aeroacoustics research in Europe: The CEAS-ASC report on 2015 highlights

    NASA Astrophysics Data System (ADS)

    Jiříček, Ondřej

    2016-10-01

    The Council of European Aerospace Societies (CEAS) Aeroacoustics Specialists Committee (ASC) supports and promotes the interests of the scientific and industrial aeroacoustics community on the European scale, and European aeronautics activities internationally. Each year, the committee highlights several of the research and development projects in Europe. This paper is the 2015 issue of this collection of Aeroacoustic Highlights, compiled from contributions submitted to the CEAS-ASC. The contributions are classified in different topics; the first categories being related to specific aeroacoustic challenges (airframe noise, fan and jet noise, helicopter noise, aircraft interior noise), while the two last sections are devoted respectively to recent improvements and emerging techniques and to general advances in aeroacoustics. Furthermore, a concise summary of the CEAS-ASC workshop "Broadband noise of rotors and airframes" held in La Rochelle, France, in September 2015 is included in this report.

  1. Comparison of NASA and contractor results from aeroacoustic tests of QCSEE OTW engine

    NASA Technical Reports Server (NTRS)

    Bloomer, H. E.; Loeffler, I. J.; Kreim, W. J.; Coats, J. W.

    1981-01-01

    The aerodynamics and acoustics of the over-the-wing (OTW) Quiet, Clean, Short Haul Experimental Engine (QCSEE) were tested. A boilerplate (nonflight weight), high-throat Mach number, acoustically treated inlet and a D-shaped OTW exhaust nozzle with variable position side doors were used. Some acoustic directivity results for the type "D" nozzle and acoustic effects of variations in the nozzle side door positions are included. It was found that the results are in agreement with those previously obtained.

  2. Description of the US Army small-scale 2-meter rotor test system

    NASA Technical Reports Server (NTRS)

    Phelps, Arthur E., III; Berry, John D.

    1987-01-01

    A small-scale powered rotor model was designed for use as a research tool in the exploratory testing of rotors and helicopter models. The model, which consists of a 29 hp rotor drive system, a four-blade fully articulated rotor, and a fuselage, was designed to be simple to operate and maintain in wind tunnels of moderate size and complexity. Two six-component strain-gauge balances are used to provide independent measurement of the rotor and fuselage aerodynamic loads. Commercially available standardized hardware and equipment were used to the maximum extent possible, and specialized parts were designed so that they could be fabricated by normal methods without using highly specialized tooling. The model was used in a hover test of three rotors having different planforms and in a forward flight investigation of a 21-percent-scale model of a U.S. Army scout helicopter equipped with a mast-mounted sight.

  3. Full-scale testing of an Ogee tip rotor. [in the Langley whirl tower

    NASA Technical Reports Server (NTRS)

    Mantay, W. R.; Campbell, R. L.; Shidler, P. A.

    1978-01-01

    Full scale tests were utilized to investigate the effect of the ogee tip on helicopter rotor acoustics, performance, and loads. Two facilities were used: the Langley whirl tower and a UH-1H helicopter. The text matrix for hover on the whirl tower involved thrust values from 0 to 44 480 N (10,000 lb) at several tip Mach numbers for both standard and Ogee rotors. The full scale testing on the UH-1H encompassed the major portion of the flight envelope for that aircraft. Both near field acoustic measurements and far field flyover data were obtained for both the ogee and standard rotors. Data analysis of the whirl tower test shows that the ogee tip does significantly diffuse the tip vortex while providing some improvement in hover performance at low and moderate thrust coefficients. Flight testing of both rotors indicates that the strong impulsive noise signature of the standard rotor can be reduced with the ogee rotor. Analysis of the spectra indicates a reduction in energy in the 250 Hz and 1000 Hz range for the ogee rotor. Forward flight performance was significantly improved with the ogee configuration for a large number of flight conditions. Further, rotor control loads were reduced through use of this advanced tip rotor.

  4. Test rig for investigations of force excited and synchrocoupling loaded rotor

    NASA Astrophysics Data System (ADS)

    Rauch, Adam

    1987-10-01

    The main topics of test rig based rotor investigations are: dynamic relations of a rotor bearing foundation system, shaft cracking, torsional simulation of rotor-generator hunting, flexural-torsional coupling of vibrations, and other areas of rotor dynamics including balancing, maintenance and modal analysis of a rotor. The present paper describes a general purpose rotor test rig capable of handling a great number of these areas. The test rig simulates a heavy, power-loaded rotor mounted on a flexible foundation. Five forms of excitation are provided: oscillating or impact torque, impact force, step force, bearing and foundation excitation. They can be combined if necessary. New research facilities offered by the rig are: external flexural force loading, driving torque loading by synchrocoupling with a full recovery of brake energy, bending release of rotor ends by pivoting, and multi-pulse impact force excitation. The most remarkable is the synchroaxle principle, called SAP for short, which is described in detail. Test rig features have been proved to be successful by research of shaft cracking and rotor-foundation vibrations. This presentation concerns only the general description, calculation and application aspects of the rig.

  5. Performance testing of a Savonius windmill rotor in shear flows

    NASA Astrophysics Data System (ADS)

    Mojola, O. O.; Onasanya, O. E.

    The effects of flow shear and/or unsteady behavior on the power generation capability of a Savonius wind turbine rotor are assessed in view of measurements conducted, both in two statistically steady shear flows and in the wind, of rotor tip speed and torque at a number of streamwise stations for each of four values of the rotor bucket overlap ratio. It is found that, even in the absence of shear, the power coefficient of a Savonius wind turbine rotor is most strongly dependent on tip speed ratio.

  6. Takeoff/approach noise for a model counterrotation propeller with a forward-swept upstream rotor

    NASA Technical Reports Server (NTRS)

    Woodward, Richard P.; Hall, David G.; Podboy, Gary G.; Jeracki, Robert J.

    1993-01-01

    A scale model of a counterrotating propeller with forward-swept blades in the forward rotor and aft-swept blades in the aft rotor (designated F39/A31) has been tested in the NASA Lewis 9- by 15-Foot Anechoic Wind Tunnel. This paper presents aeroacoustic results at a takeoff/approach condition of Mach 0.20. Laser Doppler Velocimeter results taken in a plane between the two rotors are also included to quantify the interaction flow field. The intention of the forward-swept design is to reduce the magnitude of the forward rotor tip vortex and/or wakes which impinge on the aft rotor, thus lowering the interaction tone levels.

  7. Rotor systems research aircraft risk-reduction shake test

    NASA Technical Reports Server (NTRS)

    Wellman, J. Brent

    1990-01-01

    A shake test and an extensive analysis of results were performed to evaluate the possibility of and the method for dynamically calibrating the Rotor Systems Research Aircraft (RSRA). The RSRA airframe was subjected to known vibratory loads in several degrees of freedom and the responses of many aircraft transducers were recorded. Analysis of the transducer responses using the technique of dynamic force determination showed that the RSRA, when used as a dynamic measurement system, could predict, a posteriori, an excitation force in a single axis to an accuracy of about 5 percent and sometimes better. As the analysis was broadened to include multiple degrees of freedom for the excitation force, the predictive ability of the measurement system degraded to about 20 percent, with the error occasionally reaching 100 percent. The poor performance of the measurement system is explained by the nonlinear response of the RSRA to vibratory forces and the inadequacy of the particular method used in accounting for this nonlinearity.

  8. Final project report: High energy rotor development, test and evaluation

    SciTech Connect

    1996-09-01

    Under the auspices of the {open_quotes}Government/Industry Wind Technology Applications Project{close_quotes} [{open_quotes}Letter of Interest{close_quotes} (LOI) Number RC-1-11101], Flo Wind Corp. has successfully developed, tested, and delivered a high-energy rotor upgrade candidate for their 19-meter Vertical Axis Wind Turbine. The project included the demonstration of the innovative extended height-to-diameter ratio concept, the development of a continuous span single-piece composite blade, the demonstration of a continuous blade manufacturing technique, the utilization of the Sandia National Laboratories developed SNLA 2150 natural laminar flow airfoil and the reuse of existing wind turbine and wind power plant infrastructure.

  9. Wind tunnel test of a smart rotor with individual blade twist control

    NASA Astrophysics Data System (ADS)

    Chen, Peter C.; Chopra, Inderjit

    1997-06-01

    The objective of this research is to develop a smart rotor with active control of blade twist using embedded piezoceramic elements as sensors and actuators to minimize rotor vibrations. A 1/8-th Froude-scale (dynamically-scaled) bearingless helicopter rotor model was built with banks of torsional actuators capable of manipulating blade twist at frequencies from 5 to 100 Hz. To assess the effectiveness of the torsional actuators and vibration suppression capabilities, systematic wind tunnel testing was conducted in the Glenn L. Martin Wind Tunnel. Using accelerometers embedded in the blade tip, the oscillatory blade twist response was measured. The changes in rotor vibratory loads due to piezo- induced twist were determined using a rotating hub balance located at the rotor hub. Experimental test results show that tip twist amplitudes on the order of 0.5 deg are attainable by the current actuator configurations in forward flight. Although these amplitudes were less than the target value (1 - 2 deg for complete vibration suppression control), test results show that partial vibration reduction is possible. Using open-loop phase shift control of blade twist at the first four rotor harmonics, changes in rotor thrust of up to 9% of the steady-state values were measured, resulting in up to 3 and 8% reductions in rotor pitching and rolling moments, respectively. It is expected that the hub load control authority of the smart rotor can be improved in future models with refined actuator configurations and implementation of closed-loop feedback controls.

  10. Performance testing of a Savonius windmill rotor in shear flows

    SciTech Connect

    Mojola, O.O.; Onasanya, O.E.

    1981-08-01

    The effects of flow shear and/or unsteadiness on the power-producing performance of a Savonius windmill rotor are discussed. Measurements were made, in two statistically steady shear flows and in the natural wind, of the speed, torque and (hence) power of the rotor at a number of streamwise stations for each of four values of the bucket overlap ratio. 8 refs.

  11. Boeing Smart Rotor Full-scale Wind Tunnel Test Data Report

    NASA Technical Reports Server (NTRS)

    Kottapalli, Sesi; Hagerty, Brandon; Salazar, Denise

    2016-01-01

    A full-scale helicopter smart material actuated rotor technology (SMART) rotor test was conducted in the USAF National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel at NASA Ames. The SMART rotor system is a five-bladed MD 902 bearingless rotor with active trailing-edge flaps. The flaps are actuated using piezoelectric actuators. Rotor performance, structural loads, and acoustic data were obtained over a wide range of rotor shaft angles of attack, thrust, and airspeeds. The primary test objective was to acquire unique validation data for the high-performance computing analyses developed under the Defense Advanced Research Project Agency (DARPA) Helicopter Quieting Program (HQP). Other research objectives included quantifying the ability of the on-blade flaps to achieve vibration reduction, rotor smoothing, and performance improvements. This data set of rotor performance and structural loads can be used for analytical and experimental comparison studies with other full-scale rotor systems and for analytical validation of computer simulation models. The purpose of this final data report is to document a comprehensive, highquality data set that includes only data points where the flap was actively controlled and each of the five flaps behaved in a similar manner.

  12. Analysis of the wind tunnel test of a tilt rotor power force model

    NASA Technical Reports Server (NTRS)

    Marr, R. L.; Ford, D. G.; Ferguson, S. W.

    1974-01-01

    Two series of wind tunnel tests were made to determine performance, stability and control, and rotor wake interaction on the airframe, using a one-tenth scale powered force model of a tilt rotor aircraft. Testing covered hover (IGE/OCE), helicopter, conversion, and airplane flight configurations. Forces and moments were recorded for the model from predetermined trim attitudes. Control positions were adjusted to trim flight (one-g lift, pitching moment and drag zero) within the uncorrected test data balance accuracy. Pitch and yaw sweeps were made about the trim attitudes with the control held at the trimmed settings to determine the static stability characteristics. Tail on, tail off, rotors on, and rotors off configurations were testes to determine the rotor wake effects on the empennage. Results are presented and discussed.

  13. Documentation of the Recirculation in a Closed-Chamber Rotor Hover Test

    NASA Technical Reports Server (NTRS)

    McCoy, Miranda; Wadcock, Alan J.; Young, Larry A.

    2016-01-01

    A rotor hover test was performed inside the JPL 25-foot-diameter Space Simulator. The 40-inch-diameter rotor was tested at two locations in the chamber-on the chamber centerline and 2m off-axis. The rotor was tested in both upright and inverted configurations for 500 < RPM < 2000. Fluorescent tufts were used to identify regions of recirculation. Velocities on the entrainment side of the rotor were measured. Tabulated values for the mean entrainment velocity components and the corresponding root mean square velocity fluctuations are provided. Unsteady velocity measurements provide a description of the turbulence ingested into the rotor plane and quantify the unsteady velocity field that the Mars Scout Helicopter can expect to encounter during free flight inside the Space Simulator.

  14. Helicopter blade dynamic loads measured during performance testing of two scaled rotors

    NASA Technical Reports Server (NTRS)

    Berry, John D.

    1987-01-01

    A test to determine the performance differences between the 27-percent-scale models of two rotors for the U.S. Army AH-64 helicopter was conducted in the Langley 14- by 22-Foot Subsonic Tunnel. One rotor, referred to as the baseline rotor, simulated the geometry and dynamic characteristics of the production baseline rotor, and the other rotor, referred to as the advanced rotor, was designed to have improved hover performance. During the performance test, the dynamic pitch-link forces and blade bending and torsion moments were also measured. Dynamic data from the forward flight investigation are reduced and presented. The advanced blade set was designed to have dynamic characteristics similar to those of the baseline rotor so that test conditions would not be limited by potential rotor instability and blade resonances, and so that the measured performance increments could be considered to be due purely to aerodynamic causes. Data show consistent trends with advance ratio for both blade sets with generally higher oscillatory loads occurring for the advanced blade set when compared with the baseline blade set.

  15. Testing and Life Prediction for Composite Rotor Hub Flexbeams

    NASA Technical Reports Server (NTRS)

    Murri, Gretchen B.

    2004-01-01

    A summary of several studies of delamination in tapered composite laminates with internal ply-drops is presented. Initial studies used 2D FE models to calculate interlaminar stresses at the ply-ending locations in linear tapered laminates under tension loading. Strain energy release rates for delamination in these laminates indicated that delamination would likely start at the juncture of the tapered and thin regions and grow unstably in both directions. Tests of glass/epoxy and graphite/epoxy linear tapered laminates under axial tension delaminated as predicted. Nonlinear tapered specimens were cut from a full-size helicopter rotor hub and were tested under combined constant axial tension and cyclic transverse bending loading to simulate the loading experienced by a rotorhub flexbeam in flight. For all the tested specimens, delamination began at the tip of the outermost dropped ply group and grew first toward the tapered region. A 2D FE model was created that duplicated the test flexbeam layup, geometry, and loading. Surface strains calculated by the model agreed very closely with the measured surface strains in the specimens. The delamination patterns observed in the tests were simulated in the model by releasing pairs of MPCs along those interfaces. Strain energy release rates associated with the delamination growth were calculated for several configurations and using two different FE analysis codes. Calculations from the codes agreed very closely. The strain energy release rate results were used with material characterization data to predict fatigue delamination onset lives for nonlinear tapered flexbeams with two different ply-dropping schemes. The predicted curves agreed well with the test data for each case studied.

  16. NASA rotor system research aircraft flight-test data report: Helicopter and compound configuration

    NASA Technical Reports Server (NTRS)

    Erickson, R. E.; Kufeld, R. M.; Cross, J. L.; Hodge, R. W.; Ericson, W. F.; Carter, R. D. G.

    1984-01-01

    The flight test activities of the Rotor System Research Aircraft (RSRA), NASA 740, from June 30, 1981 to August 5, 1982 are reported. Tests were conducted in both the helicopter and compound configurations. Compound tests reconfirmed the Sikorsky flight envelope except that main rotor blade bending loads reached endurance at a speed about 10 knots lower than previously. Wing incidence changes were made from 0 to 10 deg.

  17. Vibratory Loads Data from a Wind-Tunnel Test of Structurally Tailored Model Helicopter Rotors

    NASA Technical Reports Server (NTRS)

    Yeager, William T., Jr.; Hamouda, M-Nabil H.; Idol, Robert F.; Mirick, Paul H.; Singleton, Jeffrey D.; Wilbur, Matthew L.

    1991-01-01

    An experimental study was conducted in the Langley Transonic Dynamics Tunnel to investigate the use of a Bell Helicopter Textron (BHT) rotor structural tailoring concept, known as rotor nodalization, in conjunction with advanced blade aerodynamics as well as to evaluate rotor blade aerodynamic design methodologies. A 1/5-size, four-bladed bearingless hub, three sets of Mach-scaled model rotor blades were tested in forward flight from transition up to an advance ratio of 0.35. The data presented pertain only to the evaluation of the structural tailoring concept and consist of fixed-system and rotating system vibratory loads. These data will be useful for evaluating the effects of tailoring blade structural properties on fixed-system vibratory loads, as well as validating analyses used in the design of advanced rotor systems.

  18. V/STOL tilt rotor study. Volume 6: Hover, low speed and conversion tests of a tilt rotor aeroelastic model (Model 300)

    NASA Technical Reports Server (NTRS)

    Marr, R. L.; Sambell, K. W.; Neal, G. T.

    1973-01-01

    Stability and control tests of a scale model of a tilt rotor research aircraft were conducted. The characteristics of the model for hover, low speed, and conversion flight were analyzed. Hover tests were conducted in a rotor whirl cage. Helicopter and conversion tests were conducted in a low speed wind tunnel. Data obtained from the tests are presented as tables and graphs. Diagrams and illustrations of the test equipment are provided.

  19. Wind Tunnel Tests on a Different Phase Three-Stage Savonius Rotor

    NASA Astrophysics Data System (ADS)

    Hayashi, Tsutomu; Li, Yan; Hara, Yutaka

    In order to decrease the torque variation of a Savonius rotor and improve the starting characteristics, a new type of Savonius rotor, which has three stages with 120-degree bucket phase shift between the adjacent stages, has been designed and made. Wind tunnel tests make it clear that both the static and dynamic torque variations in one revolution of this three-stage rotor have been greatly smoothed in comparison with an ordinary one-stage rotor, which means the improvement of the starting characteristics. The torque characteristics of the rotors with guide vanes were also measured. The guide vanes increased the torque coefficient on the average in the low tip speed ratio but decreased the torque coefficient in high tip speed ratio. Although the present three-stage rotor needs improvement of the aspect ratio of each stage, the three-stage rotor with no guide vane had better torque characteristics than the one-stage rotor with guide vanes for tip speed ratio larger than 0.8.

  20. Investigation of UH-60A Rotor Structural Loads from Flight and Wind Tunnel Tests

    DTIC Science & Technology

    2016-05-19

    Investigation of UH-60A Rotor Structural Loads From Flight and Wind Tunnel Tests Hyeonsoo Yeo Mark Potsdam US Army Aviation Development Directorate... Aviation & Missile Research, Development & Engineering Center Research, Development, and Engineering Command Ames Research Center, Moffett Field

  1. Dynamic structural aeroelastic stability testing of the XV-15 tilt rotor research aircraft

    NASA Technical Reports Server (NTRS)

    Schroers, L. G.

    1982-01-01

    For the past 20 years, a significant effort has been made to understand and predict the structural aeroelastic stability characteristics of the tilt rotor concept. Beginning with the rotor-pylon oscillation of the XV-3 aircraft, the problem was identified and then subjected to a series of theoretical studies, plus model and full-scale wind tunnel tests. From this data base, methods were developed to predict the structural aeroelastic stability characteristics of the XV-15 Tilt Rotor Research Aircraft. The predicted aeroelastic characteristics are examined in light of the major parameters effecting rotor-pylon-wing stability. Flight test techniques used to obtain XV-15 aeroelastic stability are described. Flight test results are summarized and compared to the predicted values. Wind tunnel results are compared to flight test results and correlated with predicted values.

  2. Full-scale wind-tunnel test of the aeroelastic stability of a bearingless main rotor

    NASA Technical Reports Server (NTRS)

    Warmbrodt, W.; Mccloud, J., III; Sheffler, M.; Staley, J.

    1981-01-01

    The rotor studied in the wind tunnel had previously been flight tested on a BO-105 helicopter. The investigation was conducted to determine the rotor's aeroelastic stability characteristics in hover and at airspeeds up to 143 knots. These characteristics are compared with those obtained from whirl-tower and flight tests and predictions from a digital computer simulation. It was found that the rotor was stable for all conditions tested. At constant tip speed, shaft angle, and airspeed, stability increases with blade collective pitch setting. No significant change in system damping occurred that was attributable to frequency coalescence between the rotor inplane regressing mode and the support modes. Stability levels determined in the wind tunnel were of the same magnitude and yielded the same trends as data obtained from whirl-tower and flight tests.

  3. Small-scale rotor test rig capabilities for testing vibration alleviation algorithms

    NASA Technical Reports Server (NTRS)

    Jacklin, Stephen A.; Leyland, Jane Anne

    1987-01-01

    A test was conducted to assess the capabilities of a small scale rotor test rig for implementing higher harmonic control and stability augmentation algorithms. The test rig uses three high speed actuators to excite the swashplate over a range of frequencies. The actuator position signals were monitored to measure the response amplitudes at several frequencies. The ratio of response amplitude to excitation amplitude was plotted as a function of frequency. In addition to actuator performance, acceleration from six accelerometers placed on the test rig was monitored to determine whether a linear relationship exists between the harmonics of N/Rev control input and the least square error (LSE) identification technique was used to identify local and global transfer matrices for two rotor speeds at two batch sizes each. It was determined that the multicyclic control computer system interfaced very well with the rotor system and kept track of the input accelerometer signals and their phase angles. However, the current high speed actuators were found to be incapable of providing sufficient control authority at the higher excitation frequencies.

  4. Tests of Full-Scale Helicopter Rotors at High Advancing Tip Mach Numbers and Advance Ratios

    DTIC Science & Technology

    2015-05-01

    NASA/TM—2015–218813 Tests of Full -Scale Helicopter Rotors at High Advancing Tip Mach Numbers and Advance Ratios James C. Biggers and...Information Desk Mail Stop 148 NASA Langley Research Center Hampton, VA 23681-2199 This page is required and contains approved text that cannot be...changed. NASA/TM—2015–218813 Tests of Full -Scale Helicopter Rotors at High Advancing Tip Mach Numbers and Advance Ratios James C

  5. On determining the acoustic properties of main helicopter rotor models on an open test bench

    NASA Astrophysics Data System (ADS)

    Kop'ev, V. F.; Zaitsev, M. Yu.; Ostrikov, N. N.; Denisov, S. L.; Makashov, S. Yu.; Anikin, V. A.; Gromov, V. V.

    2016-11-01

    The paper presents the results of experimental studies on developing a technique to determine the acoustic properties of models of main helicopter rotors on an open test bench. The method of maximum length sequences is used to choose the optimum arrangement of microphones for an open test bench that would minimize the influence of parasitic echo. The results of processing the data of an acoustic experiment with a model rotor are detailed.

  6. Performance and loads data from a wind tunnel test of a full-scale, coaxial, hingeless rotor helicopter

    NASA Technical Reports Server (NTRS)

    Felker, F. F., III

    1981-01-01

    A full-scale XH-59A advancing blade concept helicopter was tested in Ames Research Center's 40 by 80 foot wind tunnel. The helicopter was tested with the rotor on and off, rotor hub fairings on and off, interrotor shaft fairing on and off, rotor instrumentation module on and off, and auxiliary propulsion thrust on and off. An advance ratio range of 0.25 and 0.45 with the rotor on and from 60 to 180 knots with the rotor off was investigated. Data on aerodynamic forces and moments, rotor loads, rotor control positions and vibration for the XH-59A as well as the aerodynamic performance of the isolated rotor are presented.

  7. Development and Hover Testing of the Active Elevon Rotor

    DTIC Science & Technology

    2012-05-01

    IBC) of rotor pitch links (Refs. 2 and 3), and various on-blade controls – such as active flaps (Refs. 4, 5, 6, 7, and 8) and active twist (Ref. 9...It is a fully articulated, four bladed hub with coincident flap and lead-lag hinges with a 2.97 inch hinge offset. The pitch axis is further...frequency of the rotor regressing lag mode. The hub has a large deflection range, ±9o in lead-lag, ±24o in pitch and +11 to -8o in flapping . Hub

  8. Aeroacoustic Prediction Codes

    NASA Technical Reports Server (NTRS)

    Gliebe, P; Mani, R.; Shin, H.; Mitchell, B.; Ashford, G.; Salamah, S.; Connell, S.; Huff, Dennis (Technical Monitor)

    2000-01-01

    This report describes work performed on Contract NAS3-27720AoI 13 as part of the NASA Advanced Subsonic Transport (AST) Noise Reduction Technology effort. Computer codes were developed to provide quantitative prediction, design, and analysis capability for several aircraft engine noise sources. The objective was to provide improved, physics-based tools for exploration of noise-reduction concepts and understanding of experimental results. Methods and codes focused on fan broadband and 'buzz saw' noise and on low-emissions combustor noise and compliment work done by other contractors under the NASA AST program to develop methods and codes for fan harmonic tone noise and jet noise. The methods and codes developed and reported herein employ a wide range of approaches, from the strictly empirical to the completely computational, with some being semiempirical analytical, and/or analytical/computational. Emphasis was on capturing the essential physics while still considering method or code utility as a practical design and analysis tool for everyday engineering use. Codes and prediction models were developed for: (1) an improved empirical correlation model for fan rotor exit flow mean and turbulence properties, for use in predicting broadband noise generated by rotor exit flow turbulence interaction with downstream stator vanes: (2) fan broadband noise models for rotor and stator/turbulence interaction sources including 3D effects, noncompact-source effects. directivity modeling, and extensions to the rotor supersonic tip-speed regime; (3) fan multiple-pure-tone in-duct sound pressure prediction methodology based on computational fluid dynamics (CFD) analysis; and (4) low-emissions combustor prediction methodology and computer code based on CFD and actuator disk theory. In addition. the relative importance of dipole and quadrupole source mechanisms was studied using direct CFD source computation for a simple cascadeigust interaction problem, and an empirical combustor

  9. Effect of vane opening on aerodynamic performance of the ram-rotor test system

    NASA Astrophysics Data System (ADS)

    Han, Ji-ang; Guan, Jian; Zhong, Jingjun; Yuan, Chenguang

    2016-06-01

    In order to research the influence of adjustable vane on the aerodynamic performance of the ram-rotor test system, FLUENT software has been adopted to simulate the flow passage of the ram-rotor test system numerically. The vane opening is controlled by changing the stagger angle of the vane blades. Results show that flow uniformity of vane outlet is influenced by the vane openings, which has an impact on the aerodynamic loss to some extent. Total pressure ratio, adiabatic efficiency and mass flow rate can be regulated by different openings of the vane. Compared with -8° vane opening, top efficiency of the ram-rotor increases by about 13.8% at +6° opening. And total pressure ratio drops by 5.87%. The rising opening increases the relative Mach number at inlet of the ram-rotor and weakens the intensity of the tip clearance leakage, which comes to a decreasing aerodynamic loss.

  10. Design, fabrication, and test of a composite material wind turbine rotor blade

    NASA Technical Reports Server (NTRS)

    Griffee, D. G., Jr.; Gustafson, R. E.; More, E. R.

    1977-01-01

    The aerodynamic design, structural design, fabrication, and structural testing is described for a 60 foot long filament wound, fiberglass/epoxy resin matrix wind turbine rotor blade for a 125 foot diameter, 100 kW wind energy conversion system. One blade was fabricated which met all aerodynamic shape requirements and was structurally capable of operating under all specified design conditions. The feasibility of filament winding large rotor blades was demonstrated.

  11. Wind Tunnel Interference Effects on Tilt Rotor Testing Using Computational Fluid Dynamics

    NASA Technical Reports Server (NTRS)

    Koning, Witold J. F.

    2015-01-01

    Experimental techniques to measure rotorcraft aerodynamic performance are widely used. However, most of them are either unable to capture interference effects from bodies, or require an extremely large computational budget. The objective of the present research is to develop an XV-15 Tilt Rotor Research Aircraft rotor model for investigation of wind tunnel wall interference using a novel Computational Fluid Dynamics (CFD) solver for rotorcraft, RotCFD. In RotCFD, a mid-fidelity URANS solver is used with an incompressible flow model and a realizable k-e turbulence model. The rotor is, however, not modeled using a computationally expensive, unsteady viscous body-fitted grid, but is instead modeled using a blade element model with a momentum source approach. Various flight modes of the XV-15 isolated rotor, including hover, tilt and airplane mode, have been simulated and correlated to existing experimental and theoretical data. The rotor model is subsequently used for wind tunnel wall interference simulations in the National Full-Scale Aerodynamics Complex (NFAC) at NASA Ames Research Center in California. The results from the validation of the isolated rotor performance showed good correlation with experimental and theoretical data. The results were on par with known theoretical analyses. In RotCFD the setup, grid generation and running of cases is faster than many CFD codes, which makes it a useful engineering tool. Performance predictions need not be as accurate as high-fidelity CFD codes, as long as wall effects can be properly simulated. For both test sections of the NFAC wall interference was examined by simulating the XV-15 rotor in the test section of the wind tunnel and with an identical grid but extended boundaries in free field. Both cases were also examined with an isolated rotor or with the rotor mounted on the modeled geometry of the Tiltrotor Test Rig (TTR). A 'quasi linear trim' was used to trim the thrust for the rotor to compare the power as a unique

  12. Shake Test Results and Dynamic Calibration Efforts for the Large Rotor Test Apparatus

    NASA Technical Reports Server (NTRS)

    Russell, Carl R.

    2014-01-01

    Prior to the full-scale wind tunnel test of the UH-60A Airloads rotor, a shake test was completed on the Large Rotor Test Apparatus. The goal of the shake test was to characterize the oscillatory response of the test rig and provide a dynamic calibration of the balance to accurately measure vibratory hub loads. This paper provides a summary of the shake test results, including balance, shaft bending gauge, and accelerometer measurements. Sensitivity to hub mass and angle of attack were investigated during the shake test. Hub mass was found to have an important impact on the vibratory forces and moments measured at the balance, especially near the UH-60A 4/rev frequency. Comparisons were made between the accelerometer data and an existing finite-element model, showing agreement on mode shapes, but not on natural frequencies. Finally, the results of a simple dynamic calibration are presented, showing the effects of changes in hub mass. The results show that the shake test data can be used to correct in-plane loads measurements up to 10 Hz and normal loads up to 30 Hz.

  13. Fatigue life on a full scale test rig: Forged versus cast wind turbine rotor shafts

    NASA Astrophysics Data System (ADS)

    Herrmann, J.; Rauert, T.; Dalhoff, P.; Sander, M.

    2016-09-01

    To reduce uncertainties associated with the fatigue life of the highly safety relevant rotor shaft and also to review today's design practice, the fatigue behaviour will be tested on a full scale test rig. Until now tests on full scale wind turbine parts are not common. Therefore, a general lack of experience on how to perform accelerated life time tests for those components exists. To clarify how to transfer real conditions to the test environment, the arrangements and deviations for the upcoming experimental test are discussed in detail. In order to complete investigations of weight saving potentials, next to getting a better comprehension of the fatigue behaviour by executing a full scale test, a further outcome are suggestions for the usage of cast and forged materials regarding the fatigue and the remaining life of the rotor shaft. It is shown, that it is worthwhile to think about a material exchange for the forged rotor shaft.

  14. User's manual for UCAP: Unified Counter-Rotation Aero-Acoustics Program

    NASA Technical Reports Server (NTRS)

    Culver, E. M.; Mccolgan, C. J.

    1993-01-01

    This is the user's manual for the Unified Counter-rotation Aeroacoustics Program (UCAP), the counter-rotation derivative of the UAAP (Unified Aero-Acoustic Program). The purpose of this program is to predict steady and unsteady air loading on the blades and the noise produced by a counter-rotation Prop-Fan. The aerodynamic method is based on linear potential theory with corrections for nonlinearity associated with axial flux induction, vortex lift on the blades, and rotor-to-rotor interference. The theory for acoustics and the theory for individual blade loading and wakes are derived in Unified Aeroacoustics Analysis for High Speed Turboprop Aerodynamics and Noise, Volume 1 (NASA CR-4329). This user's manual also includes a brief explanation of the theory used for the modelling of counter-rotation.

  15. Results of Aero/Acoustic Tests and Analytical Studies of a Two-Dimensional Eight-Lobe Mixer-Ejector Exhaust Nozzle at Takeoff Conditions

    NASA Technical Reports Server (NTRS)

    Harrington, Douglas (Technical Monitor); Schweiger, P.; Stern, A.; Gamble, E.; Barber, T.; Chiappetta, L.; LaBarre, R.; Salikuddin, M.; Shin, H.; Majjigi, R.

    2005-01-01

    Hot flow aero-acoustic tests were conducted with Pratt & Whitney's High-Speed Civil Transport (HSCT) Mixer-Ejector Exhaust Nozzles by General Electric Aircraft Engines (GEAE) in the GEAE Anechoic Freejet Noise Facility (Cell 41) located in Evendale, Ohio. The tests evaluated the impact of various geometric and design parameters on the noise generated by a two-dimensional (2-D) shrouded, 8-lobed, mixer-ejector exhaust nozzle. The shrouded mixer-ejector provides noise suppression by mixing relatively low energy ambient air with the hot, high-speed primary exhaust jet. Additional attenuation was obtained by lining the shroud internal walls with acoustic panels, which absorb acoustic energy generated during the mixing process. Two mixer designs were investigated, the high mixing "vortical" and aligned flow "axial", along with variations in the shroud internal mixing area ratios and shroud length. The shrouds were tested as hardwall or lined with acoustic panels packed with a bulk absorber. A total of 21 model configurations at 1:11.47 scale were tested. The models were tested over a range of primary nozzle pressure ratios and primary exhaust temperatures representative of typical HSCT aero thermodynamic cycles. Static as well as flight simulated data were acquired during testing. A round convergent unshrouded nozzle was tested to provide an acoustic baseline for comparison to the test configurations. Comparisons were made to previous test results obtained with this hardware at NASA Glenn's 9- by 15-foot low-speed wind tunnel (LSWT). Laser velocimetry was used to investigate external as well as ejector internal velocity profiles for comparison to computational predictions. Ejector interior wall static pressure data were also obtained. A significant reduction in exhaust system noise was demonstrated with the 2-D shrouded nozzle designs.

  16. Development and whirl tower test of the SMART active flap rotor

    NASA Astrophysics Data System (ADS)

    Straub, Friedrich K.; Kennedy, Dennis K.; Stemple, Alan D.; Anand, V. R.; Birchette, Terry S.

    2004-07-01

    A full scale Smart Material Actuated Rotor Technology (SMART) system with piezoelectric actuated blade flaps was developed and whirl tower tested. The development effort included design, fabrication, and component testing of rotor blades, trailing edge flaps, piezoelectric actuators, switching power amplifiers, and the data/power system. Simulations and model scale wind tunnel tests have shown that this system can provide 80% vibration reduction, 10dB noise reduction for a helicopter passing overhead, and substantial aerodynamic performance gains. Whirl tower testing of the 34-foot diameter rotor demonstrated the functionality, robustness, and required authority of the active flap system. The program involved extensive development work and risk reduction tests which resulted in a robust, high performance actuator and a tightly integrated actuator, flap, and blade system. The actuator demonstrated excellent performance during bench testing and has accumulated over 60 million cycles under a spectrum of loading conditions. The flight worthy active flap rotor blades were based on a modified design of the FAA certified MD900 Explorer production rotor blade. Whirl tower testing was conducted with full rotor instrumentation and a 5-component balance. The rotor was tested for 13 hours under a range of conditions, including 7 hours of flap operation. Flap inputs included open loop static and dynamic commands. The flaps showed excellent authority with oscillatory thrust greater than 10% of the steady baseline thrust. Various flap actuation frequency sweeps were run to investigate the dynamics of the rotor and the flap system. Limited closed loop tests used hub accelerations and hub loads for feedback. Proving the integration, robust operation, and authority of the flap system were the key objectives met by the whirl tower test. This success depended on tailoring the piezoelectric materials and actuator to the application and meeting actuator/blade integration requirements

  17. Blade Displacement Measurement Technique Applied to a Full-Scale Rotor Test

    NASA Technical Reports Server (NTRS)

    Abrego, Anita I.; Olson, Lawrence E.; Romander, Ethan A.; Barrows, Danny A.; Burner, Alpheus W.

    2012-01-01

    Blade displacement measurements using multi-camera photogrammetry were acquired during the full-scale wind tunnel test of the UH-60A Airloads rotor, conducted in the National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel. The objectives were to measure the blade displacement and deformation of the four rotor blades as they rotated through the entire rotor azimuth. These measurements are expected to provide a unique dataset to aid in the development and validation of rotorcraft prediction techniques. They are used to resolve the blade shape and position, including pitch, flap, lag and elastic deformation. Photogrammetric data encompass advance ratios from 0.15 to slowed rotor simulations of 1.0, thrust coefficient to rotor solidity ratios from 0.01 to 0.13, and rotor shaft angles from -10.0 to 8.0 degrees. An overview of the blade displacement measurement methodology and system development, descriptions of image processing, uncertainty considerations, preliminary results covering static and moderate advance ratio test conditions and future considerations are presented. Comparisons of experimental and computational results for a moderate advance ratio forward flight condition show good trend agreements, but also indicate significant mean discrepancies in lag and elastic twist. Blade displacement pitch measurements agree well with both the wind tunnel commanded and measured values.

  18. In-process, non-destructive, dynamic testing of high-speed polymer composite rotors

    NASA Astrophysics Data System (ADS)

    Kuschmierz, Robert; Filippatos, Angelos; Günther, Philipp; Langkamp, Albert; Hufenbach, Werner; Czarske, Jürgen; Fischer, Andreas

    2015-03-01

    Polymer composite rotors are lightweight and offer great perspectives in high-speed applications such as turbo machinery. Currently, novel rotor structures and materials are investigated for the purpose of increasing machine efficiency and lifetime, as well as allowing for higher dynamic loads. However, due to the complexity of the composite materials an in-process measurement system is required. This allows for monitoring the evolution of damages under dynamic loads, for testing and predicting the structural integrity of composite rotors in process. In rotor design, it can be used for calibrating and improving models, simulating the dynamic behaviour of polymer composite rotors. The measurement system is to work non-invasive, offer micron uncertainty, as well as a high measurement rate of several tens of kHz. Furthermore, it must be applicable at high surface speeds and under technical vacuum. In order to fulfil these demands a novel laser distance measurement system was developed. It provides the angle resolved measurement of the biaxial deformation of a fibre-reinforced polymer composite rotor with micron uncertainty at surface speeds of more than 300 m/s. Furthermore, a simulation procedure combining a finite element model and a damage mechanics model is applied. A comparison of the measured data and the numerically calculated data is performed to validate the simulation towards rotor expansion. This validating procedure can be used for a model calibration in the future. The simulation procedure could be used to investigate different damage-test cases of the rotor, in order to define its structural behaviour without further experiments.

  19. Wind Tunnel Interference Effects on Tilt Rotor Testing Using Computational Fluid Dynamics

    NASA Technical Reports Server (NTRS)

    Koning, Witold J. F.

    2016-01-01

    Experimental techniques to measure rotorcraft aerodynamic performance are widely used. However, most of them are either unable to capture interference effects from bodies, or require an extremely large computational budget. The objective of the present research is to develop an XV-15 Tiltrotor Research Aircraft rotor model for investigation of wind tunnel wall interference using a novel Computational Fluid Dynamics (CFD) solver for rotorcraft, RotCFD. In RotCFD, a mid-fidelity Unsteady Reynolds Averaged Navier-Stokes (URANS) solver is used with an incompressible flow model and a realizable k-e turbulence model. The rotor is, however, not modeled using a computationally expensive, unsteady viscous body-fitted grid, but is instead modeled using a blade-element model (BEM) with a momentum source approach. Various flight modes of the XV-15 isolated rotor, including hover, tilt, and airplane mode, have been simulated and correlated to existing experimental and theoretical data. The rotor model is subsequently used for wind tunnel wall interference simulations in the National Full-Scale Aerodynamics Complex (NFAC) at Ames Research Center in California. The results from the validation of the isolated rotor performance showed good correlation with experimental and theoretical data. The results were on par with known theoretical analyses. In RotCFD the setup, grid generation, and running of cases is faster than many CFD codes, which makes it a useful engineering tool. Performance predictions need not be as accurate as high-fidelity CFD codes, as long as wall effects can be properly simulated. For both test sections of the NFAC wall, interference was examined by simulating the XV-15 rotor in the test section of the wind tunnel and with an identical grid but extended boundaries in free field. Both cases were also examined with an isolated rotor or with the rotor mounted on the modeled geometry of the Tiltrotor Test Rig (TTR). A "quasi linear trim" was used to trim the thrust

  20. NASA rotor systems research aircraft: Fixed-wing configuration flight-test results

    NASA Technical Reports Server (NTRS)

    Erickson, R. E.; Cross, J. L.; Kufeld, R. M.; Acree, C. W.; Nguyen, D.; Hodge, R. W.

    1986-01-01

    The fixed-wing, airplane configuration flight-test results of the Rotor System Research Aircraft (RSRA), NASA 740, at Ames/Dryden Flight Research Center are documented. Fourteen taxi and flight tests were performed from December 1983 to October 1984. This was the first time the RSRA was flown with the main rotor removed; the tail rotor was installed. These tests confirmed that the RSRA is operable as a fixed-wing aircraft. Data were obtained for various takeoff and landing distances, control sensitivity, trim and dynamics stability characteristics, performance rotor-hub drag, and acoustics signature. Stability data were obtained with the rotor hub both installed and removed. The speed envelope was developed to 261 knots true airspeed (KTAS), 226 knots calibrated airspeed (KCAS) at 10,000 ft density altitude. The airplane was configured at 5 deg. wing incidence with 5 deg. wing flaps as a normal configuration. Level-flight data were acquired at 167 KCAS for wing incidence from 0 to 10 deg. Step inputs and doublet inputs of various magnitudes were utilized to acquire dynamic stability and control sensitivity data. Sine-wave inputs of constantly increasing frequency were used to generate parameter identification data. The maximum load factor attained was 2.34 g at 206 KCAS.

  1. Tower and rotor blade vibration test results for a 100-kilowatt wind turbine

    NASA Technical Reports Server (NTRS)

    Linscott, B. S.; Shapton, W. R.; Brown, D.

    1976-01-01

    The predominant natural frequencies and mode shapes for the tower and the rotor blades of the ERDA-NASA 100-kW wind turbine were determined. The tests on the tower and the blades were conducted both before and after the rotor blades and the rotating machinery were installed on top of the tower. The tower and each blade were instrumented with an accelerometer and impacted by an instrumented mass. The tower and blade structure was analyzed by means of NASTRAN, and computed values agree with the test data.

  2. Development of an Active Twist Rotor for Wind: Tunnel Testing (NLPN97-310

    NASA Technical Reports Server (NTRS)

    Cesnik, Carlos E. S.; Shin, SangJoon; Hagood, Nesbitt W., IV

    1998-01-01

    The development of the Active Twist Rotor prototype blade for hub vibration and noise reduction studies is presented in this report. Details of the modeling, design, and manufacturing are explored. The rotor blade is integrally twisted by direct strain actuation. This is accomplished by distributing embedded piezoelectric fiber composites along the span of the blade. The development of the analysis framework for this type of active blade is presented. The requirements for the prototype blade, along with the final design results are also presented. A detail discussion on the manufacturing aspects of the prototype blade is described. Experimental structural characteristics of the prototype blade compare well with design goals, and preliminary bench actuation tests show lower performance than originally predicted. Electrical difficulties with the actuators are also discussed. The presented prototype blade is leading to a complete fully articulated four-blade active twist rotor system for future wind tunnel tests.

  3. Aeromechanical stability of a hingeless rotor in hover and forward flight: Analysis and wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Yeager, W. T., Jr.; Hamouda, M. N. H.; Mantay, W. R.

    1983-01-01

    A research effort of analysis and testing was conducted to investigate the ground resonance phenomenon of a soft in-plane hingeless rotor. Experimental data were obtained using a 9 ft. (2.74 m) diameter model rotor in hover and forward flight. Eight model rotor configurations were investigated. Configuration parameters included pitch flap coupling, blade sweep and droop, and precone of the blade feathering axis. An analysis based on a comprehensive analytical model of rotorcraft aerodynamics and dynamics was used. The moving block was used to experimentally determine the regressing lead lag mode damping. Good agreement was obtained between the analysis and test. Both analysis and experiment indicated ground resonance instability in hover. An outline of the analysis, a description of the experimental model and procedures, and comparison of the analytical and experimental data are presented.

  4. Large-scale aeroacoustic research feasibility and conceptual design of test-section inserts for the Ames 80- by 120-foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Soderman, Paul T.; Olsen, Larry E.

    1990-01-01

    An engineering feasibility study was made of aeroacoustic inserts designed for large-scale acoustic research on aircraft models in the 80 by 120 foot Wind Tunnel at NASA Ames Research Center. The advantages and disadvantages of likely designs were analyzed. Results indicate that the required maximum airspeed leads to the design of a particular insert. Using goals of 200, 150, and 100 knots airspeed, the analysis indicated a 30 x 60 ft open-jet test section, a 40 x 80 ft open jet test section, and a 70 x 100 ft closed test section with enhanced wall lining, respectively. The open-jet inserts would be composed of a nozzle, collector, diffuser, and acoutic wedges incorporated in the existing 80 x 120 test section. The closed test section would be composed of approximately 5 ft acoustic wedges covered by a porous plate attached to the test section walls of the existing 80 x 120. All designs would require a double row of acoustic vanes between the test section and fan drive to attenuate fan noise and, in the case of the open-jet designs, to control flow separation at the diffuser downstream end. The inserts would allow virtually anechoic acoustic studies of large helicopter models, jets, and V/STOL aircraft models in simulated flight. Model scale studies would be necessary to optimize the aerodynamic and acoustic performance of any of the designs. In all designs studied, the existing structure would have to be reinforced. Successful development of acoustically transparent walls, though not strictly necessary to the project, would lead to a porous-wall test section that could be substituted for any of the open-jet designs, and thereby eliminate many aerodynamic and acoustic problems characteristic of open-jet shear layers. The larger size of the facility would make installation and removal of the insert components difficult. Consequently, scheduling of the existing 80 x 120 aerodynamic test section and scheduling of the open-jet test section would likely be made on an

  5. Benchmark problems in computational aeroacoustics

    NASA Technical Reports Server (NTRS)

    Porter-Locklear, Freda

    1994-01-01

    A recent directive at NASA Langley is aimed at numerically predicting principal noise sources. During my summer stay, I worked with high-order ENO code, developed by Dr. Harold Atkins, for solving the unsteady compressible Navier-Stokes equations, as it applies to computational aeroacoustics (CAA). A CAA workshop, composed of six categories of benchmark problems, has been organized to test various numerical properties of code. My task was to determine the robustness of Atkins' code for these test problems. In one category, we tested the nonlinear wave propagation of the code for the one-dimensional Euler equations, with initial pressure, density, and velocity conditions. Using freestream boundary conditions, our results were plausible. In another category, we solved the linearized two-dimensional Euler equations to test the effectiveness of radiation boundary conditions. Here we utilized MAPLE to compute eigenvalues and eigenvectors of the Jacobian given variable and flux vectors. We experienced a minor problem with inflow and outflow boundary conditions. Next, we solved the quasi one dimensional unsteady flow equations with an incoming acoustic wave of amplitude 10(exp -6). The small amplitude sound wave was incident on a convergent-divergent nozzle. After finding a steady-state solution and then marching forward, our solution indicated that after 30 periods the acoustic wave had dissipated (a period is time required for sound wave to traverse one end of nozzle to other end).

  6. Benchmark problems in computational aeroacoustics

    NASA Astrophysics Data System (ADS)

    Porter-Locklear, Freda

    1994-12-01

    A recent directive at NASA Langley is aimed at numerically predicting principal noise sources. During my summer stay, I worked with high-order ENO code, developed by Dr. Harold Atkins, for solving the unsteady compressible Navier-Stokes equations, as it applies to computational aeroacoustics (CAA). A CAA workshop, composed of six categories of benchmark problems, has been organized to test various numerical properties of code. My task was to determine the robustness of Atkins' code for these test problems. In one category, we tested the nonlinear wave propagation of the code for the one-dimensional Euler equations, with initial pressure, density, and velocity conditions. Using freestream boundary conditions, our results were plausible. In another category, we solved the linearized two-dimensional Euler equations to test the effectiveness of radiation boundary conditions. Here we utilized MAPLE to compute eigenvalues and eigenvectors of the Jacobian given variable and flux vectors. We experienced a minor problem with inflow and outflow boundary conditions. Next, we solved the quasi one dimensional unsteady flow equations with an incoming acoustic wave of amplitude 10(exp -6). The small amplitude sound wave was incident on a convergent-divergent nozzle. After finding a steady-state solution and then marching forward, our solution indicated that after 30 periods the acoustic wave had dissipated (a period is time required for sound wave to traverse one end of nozzle to other end).

  7. Rotating Shake Test and Modal Analysis of a Model Helicopter Rotor Blade

    NASA Technical Reports Server (NTRS)

    Wilkie, W. Keats; Mirick, Paul H.; Langston, Chester W.

    1997-01-01

    Rotating blade frequencies for a model generic helicopter rotor blade mounted on an articulated hub were experimentally determined. Testing was conducted using the Aeroelastic Rotor Experimental System (ARES) testbed in the Helicopter Hover Facility (HBF) at Langley Research Center. The measured data were compared to pretest analytical predictions of the rotating blade frequencies made using the MSC/NASTRAN finite-element computer code. The MSC/NASTRAN solution sequences used to analyze the model were modified to account for differential stiffening effects caused by the centrifugal force acting on the blade and rotating system dynamic effects. The correlation of the MSC/NASTRAN-derived frequencies with the experimental data is, in general, very good although discrepancies in the blade torsional frequency trends and magnitudes were observed. The procedures necessary to perform a rotating system modal analysis of a helicopter rotor blade with MSC/NASTRAN are outlined, and complete sample data deck listings are provided.

  8. Introduction to computational aero-acoustics

    NASA Technical Reports Server (NTRS)

    Hardin, Jay C.

    1996-01-01

    Computational aeroacoustics (CAA) is introduced, by presenting its definition, advantages, applications, and initial challenges. The effects of Mach number and Reynolds number on CAA are considered. The CAA method combines the methods of aeroacoustics and computational fluid dynamics.

  9. Computational Aeroacoustics: An Overview

    NASA Technical Reports Server (NTRS)

    Tam, Christopher K. W.

    2003-01-01

    An overview of recent advances in computational aeroacoustics (CAA) is presented. CAA algorithms must not be dispersive and dissipative. It should propagate waves supported by the Euler equations with the correct group velocities. Computation domains are inevitably finite in size. To avoid the reflection of acoustic and other outgoing waves at the boundaries of the computation domain, it is required that special boundary conditions be imposed at the boundary region. These boundary conditions either absorb all the outgoing waves without reflection or allow the waves to exit smoothly. High-order schemes, invariably, supports spurious short waves. These spurious waves tend to pollute the numerical solution. They must be selectively damped or filtered out. All these issues and relevant computation methods are briefly reviewed. Jet screech tones are known to have caused structural fatigue in military combat aircrafts. Numerical simulation of the jet screech phenomenon is presented as an example of a successful application of CAA.

  10. Development and Operation of an Automatic Rotor Trim Control System for the UH-60 Individual Blade Control Wind Tunnel Test

    NASA Technical Reports Server (NTRS)

    Theodore, Colin R.; Tischler, Mark B.

    2010-01-01

    An automatic rotor trim control system was developed and successfully used during a wind tunnel test of a full-scale UH-60 rotor system with Individual Blade Control (IBC) actuators. The trim control system allowed rotor trim to be set more quickly, precisely and repeatably than in previous wind tunnel tests. This control system also allowed the rotor trim state to be maintained during transients and drift in wind tunnel flow, and through changes in IBC actuation. The ability to maintain a consistent rotor trim state was key to quickly and accurately evaluating the effect of IBC on rotor performance, vibration, noise and loads. This paper presents details of the design and implementation of the trim control system including the rotor system hardware, trim control requirements, and trim control hardware and software implementation. Results are presented showing the effect of IBC on rotor trim and dynamic response, a validation of the rotor dynamic simulation used to calculate the initial control gains and tuning of the control system, and the overall performance of the trim control system during the wind tunnel test.

  11. Testing of a one-bladed 30-meter-diameter rotor on the DOE/NASA Mod-O wind turbine

    NASA Technical Reports Server (NTRS)

    Ensworth, C. B. F., III; Corrigan, R. D.; Berkowitz, B. M.

    1988-01-01

    Tests were conducted on the DOE/NASA Mod-O 200-kW horizontal-axis wind turbine in a one-bladed rotor configuration. The objectives of the test were to evaluate the performance, loads, and dynamic characteristics of a one-bladed rotor, and then to compare these parameters with those of an aerodynamically similar two-bladed rotor configuration. Test operations showed that this intermediate-size (15.2-m radius) one-bladed rotor configuration can be operated successfully. Test results show that the one-bladed rotor had cyclic blade loads comparable to those of a two-bladed rotor. A moderate power penalty equivalent to a reduction in windspeed of 1 m/sec occurred with the one-bladed rotor when operated at a rotor speed 50 percent higher than that of the two-bladed rotor.

  12. Teetered, tip-controlled rotor: Preliminary test results from Mod-0 100-kW experimental wind turbine

    NASA Technical Reports Server (NTRS)

    Glasgow, J. C.; Miller, D. R.

    1980-01-01

    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.

  13. Case Studies for the Statistical Design of Experiments Applied to Powered Rotor Wind Tunnel Tests

    NASA Technical Reports Server (NTRS)

    Overmeyer, Austin D.; Tanner, Philip E.; Martin, Preston B.; Commo, Sean A.

    2015-01-01

    The application of statistical Design of Experiments (DOE) to helicopter wind tunnel testing was explored during two powered rotor wind tunnel entries during the summers of 2012 and 2013. These tests were performed jointly by the U.S. Army Aviation Development Directorate Joint Research Program Office and NASA Rotary Wing Project Office, currently the Revolutionary Vertical Lift Project, at NASA Langley Research Center located in Hampton, Virginia. Both entries were conducted in the 14- by 22-Foot Subsonic Tunnel with a small portion of the overall tests devoted to developing case studies of the DOE approach as it applies to powered rotor testing. A 16-47 times reduction in the number of data points required was estimated by comparing the DOE approach to conventional testing methods. The average error for the DOE surface response model for the OH-58F test was 0.95 percent and 4.06 percent for drag and download, respectively. The DOE surface response model of the Active Flow Control test captured the drag within 4.1 percent of measured data. The operational differences between the two testing approaches are identified, but did not prevent the safe operation of the powered rotor model throughout the DOE test matrices.

  14. Measurement of the UH-60A Hub Large Rotor Test Apparatus Control System Stiffness

    NASA Technical Reports Server (NTRS)

    Kufeld, Robert M.

    2014-01-01

    This purpose of this report is to provides details of the measurement of the control system stiffness of the UH-60A rotor hub mounted on the Large Rotor Test Apparatus (UH-60A/LRTA). The UH-60A/LRTA was used in the 40- by 80-Foot Wind Tunnel to complete the full-scale wind tunnel test portion of the NASA / ARMY UH-60A Airloads Program. This report describes the LRTA control system and highlights the differences between the LRTA and UH-60A aircraft. The test hardware, test setup, and test procedures are also described. Sample results are shown, including the azimuthal variation of the measured control system stiffness for three different loadings and two different dynamic actuator settings. Finally, the azimuthal stiffness is converted to fixed system values using multi-blade transformations for input to comprehensive rotorcraft prediction codes.

  15. Eulerian laser Doppler vibrometry: Online blade damage identification on a multi-blade test rotor

    NASA Astrophysics Data System (ADS)

    Oberholster, A. J.; Heyns, P. S.

    2011-01-01

    Laser Doppler vibrometry enables the telemetry-free measurement of online turbomachinery blade vibration. Specifically, the Eulerian or fixed reference frame implementation of laser vibrometry provides a practical solution to the condition monitoring of rotating blades. The short data samples that are characteristic of this measurement approach do however negate the use of traditional frequency domain signal processing techniques. It is therefore necessary to employ techniques such as time domain analysis and non-harmonic Fourier analysis to obtain useful information from the blade vibration signatures. The latter analysis technique allows the calculation of phase angle trends which can be used as indicators of blade health deterioration, as has been shown in previous work for a single-blade rotor. This article presents the results from tests conducted on a five-blade axial-flow test rotor at different rotor speeds and measurement positions. With the aid of artificial neural networks, it is demonstrated that the parameters obtained from non-harmonic Fourier analysis and time domain signal processing on Eulerian laser Doppler vibrometry signals can successfully be used to identify and quantify blade damage from among healthy blades. It is also shown that the natural frequencies of individual blades can be approximated from the Eulerian signatures recorded during rotor run-up and run-down.

  16. Experimental Results of Performance Tests on a Four-Port Wave Rotor

    NASA Technical Reports Server (NTRS)

    Wilson, John; Welch, Gerard E.; Paxson, Daniel E.

    2007-01-01

    A series of tests has been performed on a four-port wave rotor suitable for use as a topping stage on a gas turbine engine, to measure the overall pressure ratio obtainable as a function of temperature ratio, inlet mass flow, loop flow ratio, and rotor speed. The wave rotor employed an open high pressure loop that is the high pressure inlet flow was not the air exhausted from the high pressure outlet, but was obtained from a separate heated source, although the mass flow rates of the two flows were balanced. This permitted the choice of a range of loop-flow ratios (i.e., ratio of high pressure flow to low pressure flow), as well as the possibility of examining the effect of mass flow imbalance. Imbalance could occur as a result of leakage or deliberate bleeding for cooling air. Measurements of the pressure drop in the high pressure loop were also obtained. A pressure ratio of 1.17 was obtained at a temperature ratio of 2.0, with an inlet mass flow of 0.6 lb/s. Earlier tests had given a pressure ratio of less than 1.12. The improvement was due to improved sealing between the high pressure and low pressure loops, and a modification to the movable end-wall which is provided to allow for rotor expansion.

  17. Hover Testing of the NASA/Army/MIT Active Twist Rotor Prototype Blade

    NASA Technical Reports Server (NTRS)

    Wilbur, Matthew L.; Yeager, William T., Jr.; Wilkie, W. Keats; Cesnik, Carlos E. S.; Shin, Sangloon

    2000-01-01

    Helicopter rotor individual blade control promises to provide a mechanism for increased rotor performance and reduced rotorcraft vibrations and noise. Active material methods, such as piezoelectrically actuated trailing-edge flaps and strain-induced rotor blade twisting, provide a means of accomplishing individual blade control without the need for hydraulic power in the rotating system. Recent studies have indicated that controlled strain induced blade twisting can be attained using piezoelectric active fiber composite technology. In order to validate these findings experimentally, a cooperative effort between NASA Langley Research Center, the Army Research Laboratory, and the MIT Active Materials and Structures Laboratory has been developed. As a result of this collaboration an aeroelastically-scaled active-twist model rotor blade has been designed and fabricated for testing in the heavy gas environment of the Langley Transonic Dynamics Tunnel (TDT). The results of hover tests of the active-twist prototype blade are presented in this paper. Comparisons with applicable analytical predictions of active-twist frequency response in hovering flight are also presented.

  18. Aeroacoustic Analysis of Fan Noise Reduction With Increased Bypass Nozzle Area

    NASA Technical Reports Server (NTRS)

    Woodward, Richard P.; Hughes, Christopher E.; Podboy, Gary G.

    2005-01-01

    An advanced model turbofan was tested in the NASA Glenn 9-by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) to explore far field acoustic effects of increased bypass nozzle area. This fan stage test was part of the NASA Glenn Fan Broadband Source Diagnostic Test, second entry (SDT2) which acquired aeroacoustic results over a range of test conditions. The baseline nozzle was sized to produce maximum stage performance at cruise condition. However, the wind tunnel testing is conducted near sea level condition. Therefore, in order to simulate and obtain performance at other operating conditions, two additional nozzles were designed and tested one with +5 percent increase in weight flow (+5.4 percent increase in nozzle area compared with the baseline nozzle), sized to simulate the performance at the stage design point (takeoff) condition, and the other with a +7.5 percent increase in weight flow (+10.9 percent increase in nozzle area) sized for maximum weight flow with a fixed nozzle at sea level condition. Measured acoustic benefits with increased nozzle area were very encouraging, showing overall sound power level (OAPWL) reductions of 2 or more dB while the stage thrust actually increased by 2 to 3 percent except for the most open nozzle at takeoff rotor speed where stage performance decreased. Effective perceived noise levels for a 1500 ft engine flyover and 3.35 scale factor showed a similar noise reduction of 2 or more EPNdB. Noise reductions, principally in the level of broadband noise, were observed everywhere in the far field. Laser Doppler Velocimetry measurements taken downstream of the rotor showed that the total turbulent velocity decreased with increasing nozzle flow, which may explain the reduced rotor broadband noise levels.

  19. Teetered, tip-controlled rotor - Preliminary test results from Mod-O 100-kW experimental wind turbine

    NASA Technical Reports Server (NTRS)

    Glasgow, J. C.; Miller, D. R.

    1980-01-01

    A series of tests is currently being conducted using the DOE/NASA 100 kW Experimental Wind Turbine with a two-bladed, teetered rotor with 30% span tip control. Preliminary evaluation test results indicates that the teetered rotor significantly decreases loads on the yaw drive mechanism and reduces blade cyclic flapwise bending moments by 25% at the 20% span location when compared to rigid hub rotor. The teetered hub performed well but did impact 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.

  20. Wind-Tunnel Tests of a Portion of a PV-2 Helicopter Rotor Blade

    NASA Technical Reports Server (NTRS)

    Kemp, William B., Jr.

    1945-01-01

    A portion of a PV-2 helicopter rotor blade has been tested in the 6- by 6-foot test section of the Langley stability tunnel to determine if the aerodynamic characteristics were seriously affected by cross flow or fabric distortion. The outer portion of the blade was tested as a reflection plane model pivoted about the tunnel wall to obtain various angles of cross flow over the blade. Because the tunnel wall acts as a plane of sytry, the measured aerodynamic characteristics correspond to those of an airfoil having various angles of sweepforward and sweepback. Tests were made with the vents on the lower surface open and also with the vents sealed and the internal pressure held at -20 inches of water producing an internal pressure coefficient of -1.059. The change in contour resulting from the range of internal pressures used had very little effect on the aerodynamic characteristics of the blade. The test methods were considered to simulate inadequately the flow conditions over the rotor blade because the effects of cross flow were limited to conditions corresponding to sweep of the blade. The results indicated that this type of cross flow had only minor effects on the aerodynamic characteristics of the blade. It is believed, therefore, that future tests to determine the effects on the aerodynamic characteristics of cross flow should utilize complete rotors.

  1. Turbine Rotor Disk Health Monitoring Assessment Based on Sensor Technology and Spin Tests Data

    PubMed Central

    2013-01-01

    The paper focuses on presenting data obtained from spin test experiments of a turbine engine like rotor disk and assessing their correlation to the development of a structural health monitoring and fault detection system. The data were obtained under various operating conditions such as the rotor disk being artificially induced with and without a notch and rotated at a rotational speed of up to 10,000 rpm under balanced and imbalanced state. The data collected included blade tip clearance, blade tip timing measurements, and shaft displacements. Two different sensor technologies were employed in the testing: microwave and capacitive sensors, respectively. The experimental tests were conducted at the NASA Glenn Research Center's Rotordynamics Laboratory using a high precision spin system. Disk flaw observations and related assessments from the collected data for both sensors are reported and discussed. PMID:23844396

  2. Turbine rotor disk health monitoring assessment based on sensor technology and spin tests data.

    PubMed

    Abdul-Aziz, Ali; Woike, Mark

    2013-01-01

    The paper focuses on presenting data obtained from spin test experiments of a turbine engine like rotor disk and assessing their correlation to the development of a structural health monitoring and fault detection system. The data were obtained under various operating conditions such as the rotor disk being artificially induced with and without a notch and rotated at a rotational speed of up to 10,000 rpm under balanced and imbalanced state. The data collected included blade tip clearance, blade tip timing measurements, and shaft displacements. Two different sensor technologies were employed in the testing: microwave and capacitive sensors, respectively. The experimental tests were conducted at the NASA Glenn Research Center's Rotordynamics Laboratory using a high precision spin system. Disk flaw observations and related assessments from the collected data for both sensors are reported and discussed.

  3. Aero-acoustic design and test of a multiple splitter exhaust noise suppressor for a 0.914m diameter lift fan

    NASA Technical Reports Server (NTRS)

    Stimpert, D. L.

    1973-01-01

    A lift fan exhaust suppression system to meet future VTOL aircraft noise goals was designed and tested. The test vehicle was a 1.3 pressure ratio, 36 inch (91.44 cm) diameter lift fan with two chord rotor to stator spacing. A two splitter fan exhaust suppression system thirty inches (76.2 cm) long achieved 10 PNdB exhaust suppression in the aft quadrant compared to a design value of 20 PNdB. It was found that a broadband noise floor limited the realizable suppression. An analytical investigation of broadband noise generated by flow over the treatment surfaces provided very good agreement with the measured suppression levels and noise floor sound power levels. A fan thrust decrement of 22% was measured for the fully suppressed configuration of which 11.1% was attributed to the exhaust suppression hardware.

  4. KSC VAB Aeroacoustic Hazard Assessment

    NASA Technical Reports Server (NTRS)

    Oliveira, Justin M.; Yedo, Sabrina; Campbell, Michael D.; Atkinson, Joseph P.

    2010-01-01

    NASA Kennedy Space Center (KSC) carried out an analysis of the effects of aeroacoustics produced by stationary solid rocket motors in processing areas at KSC. In the current paper, attention is directed toward the acoustic effects of a motor burning within the Vehicle Assembly Building (VAB). The analysis was carried out with support from ASRC Aerospace who modeled transmission effects into surrounding facilities. Calculations were done using semi-analytical models for both aeroacoustics and transmission. From the results it was concluded that acoustic hazards in proximity to the source of ignition and plume can be severe; acoustic hazards in the far-field are significantly lower.

  5. Highly Accurate Schemes for Wave Propagation Systems: Application in Aeroacoustics

    SciTech Connect

    Bartoli, Nathalie; Mazet, Pierre-Alain; Mouysset, Vincent; Rogier, Francois

    2010-09-30

    The Discontinuous Galerkin (DG) method is considered for computational aeroacoustic. A software has been developed to make it possible to test a large variety of configurations (non-conform grid, variable polynomial order). To deal with instationary phenomena involved by some shear flows, a compromise between time computation and accuracy is deduced from some numerical experiments.

  6. Development of Test Facility for Structural Evaluation of Composite Rotor Hubs

    SciTech Connect

    Dorrell, Larry; Roach, Dennis

    1999-07-20

    The rotorcraft industry is constantly evaluating new types of lightweight composite materials that not only enhance the safety and reliability of rotor components, but also improve performance and extend operating life as well. The tests required for these evaluations are typically quite complex requiring massive test fixtures, in many cases, along with multiple actuators for loading test articles at various points simultaneously. This paper discusses the background for development of the facility, as well as hardware and overall system design and implementation. Additional topics that are covered include data acquisition, implementation of nondestructive inspection techniques during the test process, and some results from the initial test series performed in the facility.

  7. Aeroacoustics research in Europe: The CEAS-ASC report on 2010 highlights

    NASA Astrophysics Data System (ADS)

    Balázs Nagy, Attila

    2011-10-01

    The Council of European Aerospace Societies (CEAS) Aeroacoustics Specialists Committee (ASC) supports and promotes the interests of the scientific and industrial aeroacoustics community on an European scale and European aeronautics activities internationally. In this context, "aeroacoustics" encompasses all aerospace acoustics and related areas. Each year the committee highlights some of the research and development projects in Europe. This paper is a report on highlights of aeroacoustics research in Europe in 2010, compiled from information provided to the ASC of the CEAS. At the end of 2010, project X-NOISE EV of the Seventh Framework Programme of the European Commission has been launched as a continuation of the X-Noise series, with objectives of reducing aircraft noise and reaching the goal set by the ACARE 2020 Vision. Some contributions submitted to the editor summarizes selected findings from European projects launched before or concluded in 2010, while other articles cover issues supported by national associations or by industries. Furthermore, a concise summary of the workshop on "Aeroacoustics of High-Speed Aircraft Propellers and Open Rotors" held in Warsaw in October is included in this report. Enquiries concerning all contributions should be addressed to the authors who are given at the end of each subsection.

  8. Measurement of rotor system dynamic stiffness by perturbation testing

    NASA Technical Reports Server (NTRS)

    Bently, D. E.; Muszynska, A.

    1985-01-01

    Specific aspects of the application of Modal Analysis to rotating machines are discussed. For lowest mode analysis, the circular force perturbation testing gives the best results. Examples of application are presented.

  9. Application of system identification to analytic rotor modeling from simulated and wind tunnel dynamic test data, part 2

    NASA Technical Reports Server (NTRS)

    Hohenemser, K. H.; Banerjee, D.

    1977-01-01

    An introduction to aircraft state and parameter identification methods is presented. A simplified form of the maximum likelihood method is selected to extract analytical aeroelastic rotor models from simulated and dynamic wind tunnel test results for accelerated cyclic pitch stirring excitation. The dynamic inflow characteristics for forward flight conditions from the blade flapping responses without direct inflow measurements were examined. The rotor blades are essentially rigid for inplane bending and for torsion within the frequency range of study, but flexible in out-of-plane bending. Reverse flow effects are considered for high rotor advance ratios. Two inflow models are studied; the first is based on an equivalent blade Lock number, the second is based on a time delayed momentum inflow. In addition to the inflow parameters, basic rotor parameters like the blade natural frequency and the actual blade Lock number are identified together with measurement bias values. The effect of the theoretical dynamic inflow on the rotor eigenvalues is evaluated.

  10. Finding the ideal strategy: Full-scale fatigue testing of wind turbine rotor shafts

    NASA Astrophysics Data System (ADS)

    Rauert, T.; Herrmann, J.; Dalhoff, P.; Sander, M.

    2016-09-01

    For the purpose of a light weight design of rotor shafts, fatigue testing is necessary. Since full-scale fatigue tests of these large components are time consuming, costly and have not been done before, much effort has to be put into the implementation of a suitable test strategy. The paper presents the boundary conditions that have to be considered to determine the finite life regime of the component S/N-curve. A statistical simulation shows how much the derived S/N-curve is influenced by the specific test procedure.

  11. Analysis of the free-tip rotor wind-tunnel test results

    NASA Technical Reports Server (NTRS)

    Stroub, Robert H.

    1985-01-01

    The results from a wind tunnel test of a small scale free-tip rotor are analyzed. The free-tip rotor has blade tips that are free to weathervane into the tip's relative wind, thus producing a more uniform lift around the azimuth. The free-tip assembly, which includes the controller, functioned flawlessly throughout the test. In a test of the free-tip's response after passing through a vertical air jet, the tip pitched freely and in a controlled manner. Analysis of the tip's response characteristics showed the free-tip system's damped natural frequency to be 5.2 per rev. Tip pitch angle responses to the local airstream are presented for an advance-ratio range of 0.1 to 0.397 and for a solidity weighted rotor lift coefficient range of 0.038 to 0.092. Harmonic analysis of the responses showed a dominance by the first harmonic. As a result of the tip being free, forward flight power requirements were reduced by 8% or more. More power reduction was recorded for high thrust conditions. In addition to the power reduction, flatwise blade bending moments were reduced by as much as 30% at the inboard blade stations.

  12. March 1971 wind tunnel tests of the Dorand DH 2011 jet flap rotor, volume 1

    NASA Technical Reports Server (NTRS)

    Kretz, M.; Aubrun, J.; Larche, M.

    1973-01-01

    The results of wind tunnel tests, second series of tests performed in the NASA Ames 40 x 80 foot wind tunnel, of the DH 2011 jet-flap rotor are presented and analyzed. The tests have been focused on multicyclic effects and the capability of this rotor to reduce the vibratory loads and stresses in the blades. The reductions of the vibrations and stresses at tip speed ratio of 0.4 have attained 50%. The theory shows further reductions possible, reaching 80%. The results show that the performance characteristics after the modifications introduced since 1965 remained unchanged. The domain of investigation has been enlarged to include the tip speed ratios of 0.6 and 0.7. To analyze the complex aeroelastic phenomena a new analytical technique has been utilized to represent the mathematical model of the rotor. This technique, based on transfer matrices and transfer functions, appears very simple and it is believed that this analysis is applicable to many kinds of investigations involving large numbers of variables.

  13. The NASA Environmentally Responsible Aviation Project/General Electric Open Rotor Test Campaign

    NASA Technical Reports Server (NTRS)

    Van Zante, Dale

    2013-01-01

    The Open Rotor is a modern version of the UnDucted Fan (UDF) that was flight tested in the late 1980's through a partnership between NASA and General Electric (GE). Tests were conducted in the 9'x15' Low Speed Wind Tunnel and the 8'x6' Supersonic Wind Tunnel starting in late 2009 and completed in early 2012. Aerodynamic and acoustic data were obtained for takeoff, approach and cruise simulations. GE was the primary partner, but other organizations were involved such as Boeing and Airbus who provided additional hardware for fuselage simulations. This test campaign provided the acoustic and performance characteristics for modern open rotor blades designs." NASA and GE conducted joint systems analysis to evaluate how well new blade designs would perform on a B737 class aircraft, and compared the results to an advanced higher bypass ratio turbofan." Acoustic shielding experiments were performed at NASA GRC and Boeing LSAF facilities to provide data for noise estimates of unconventional aircraft configurations with Open Rotor propulsion systems." The work was sponsored by NASA's aeronautics programs, including the Subsonic Fixed Wing (SFW) and the Environmentally Responsible Aviation (ERA) projects."

  14. Active Flap Control of the SMART Rotor for Vibration Reduction

    NASA Technical Reports Server (NTRS)

    Hall, Steven R.; Anand, R. Vaidyanathan; Straub, Friedrich K.; Lau, Benton H.

    2009-01-01

    Active control methodologies were applied to a full-scale active flap rotor obtained during a joint Boeing/ DARPA/NASA/Army test in the Air Force National Full-Scale Aerodynamic Complex 40- by 80-foot anechoic wind tunnel. The active flap rotor is a full-scale MD 900 helicopter main rotor with each of its five blades modified to include an on-blade piezoelectric actuator-driven flap with a span of 18% of radius, 25% of chord, and located at 83% radius. Vibration control demonstrated the potential of active flaps for effective control of vibratory loads, especially normal force loads. Active control of normal force vibratory loads using active flaps and a continuous-time higher harmonic control algorithm was very effective, reducing harmonic (1-5P) normal force vibratory loads by 95% in both cruise and approach conditions. Control of vibratory roll and pitch moments was also demonstrated, although moment control was less effective than normal force control. Finally, active control was used to precisely control blade flap position for correlation with pretest predictions of rotor aeroacoustics. Flap displacements were commanded to follow specific harmonic profiles of 2 deg or more in amplitude, and the flap deflection errors obtained were less than 0.2 deg r.m.s.

  15. Nonlinear Whirl Response of a High-Speed Seal Test Rotor With Marginal and Extended Squeeze-Film Dampers

    NASA Technical Reports Server (NTRS)

    Proctor, Margaret P.; Gunter, Edgar J.

    2005-01-01

    Synchronous and nonsynchronous whirl response analysis of a double overhung, high-speed seal test rotor with ball bearings supported in 5.84- and 12.7-mm-long, un-centered squeeze-film oil dampers is presented. Test performance with the original damper of length 5.84 mm was marginal, with nonsynchronous whirling at the overhung seal test disk and high amplitude synchronous response above 32,000 rpm near the drive spline section occurring. A system critical speed analysis of the drive system and the high-speed seal test rotor indicated that the first two critical speeds are associated with the seal test rotor. Nonlinear synchronous unbalance and time transient whirl studies were conducted on the seal test rotor with the original and extended damper lengths. With the original damper design, the nonlinear synchronous response showed that unbalance could cause damper lockup at 33,000 rpm. Alford cross-coupling forces were also included at the overhung seal test disk for the whirl analysis. Sub-synchronous whirling at the seal test disk was observed in the nonlinear time transient analysis. With the extended damper length of 12.7 mm, the sub-synchronous motion was eliminated and the rotor unbalance response was acceptable to 45,000 rpm with moderate rotor unbalance. However, with high rotor unbalance, damper lockup could still occur at 33,000 rpm, even with the extended squeeze-film dampers. Therefore, the test rotor must be reasonably balanced in order for the un-centered dampers to be effective.

  16. Parametric Blade Study Test Report Rotor Configuration. Number 4

    DTIC Science & Technology

    1988-11-01

    line, the throttle setting (with four throttle cover plates installed) which induced stall was determined; several throttle settings in the operating...was 88.5% (design efficiency was 85.4%) and stage pressure ratio was 1.92 (design stage pressure ratio was 1.92). The compressor was throttled to stall ...full open throttle to near stall (the last test point plotted on each speed line represents the highest throttle setting that could be maintained without

  17. Dynamic Open-Rotor Composite Shield Impact Test Report

    NASA Technical Reports Server (NTRS)

    Seng, Silvia; Frankenberger, Charles; Ruggeri, Charles R.; Revilock, Duane M.; Pereira, J. Michael; Carney, Kelly S.; Emmerling, William C.

    2015-01-01

    The Federal Aviation Administration (FAA) is working with the European Aviation Safety Agency to determine the certification base for proposed new engines that would not have a containment structure on large commercial aircraft. Equivalent safety to the current fleet is desired by the regulators, which means that loss of a single fan blade will not cause hazard to the aircraft. NASA Glenn and Naval Air Warfare Center (NAWC) China Lake collaborated with the FAA Aircraft Catastrophic Failure Prevention Program to design and test a shield that would protect the aircraft passengers and critical systems from a released blade that could impact the fuselage. This report documents the live-fire test from a full-scale rig at NAWC China Lake. NASA provided manpower and photogrammetry expertise to document the impact and damage to the shields. The test was successful: the blade was stopped from penetrating the shield, which validates the design analysis method and the parameters used in the analysis. Additional work is required to implement the shielding into the aircraft.

  18. Aeroacoustic research: An Army perspective

    NASA Technical Reports Server (NTRS)

    Morse, H. A.; Schmitz, F. H.

    1978-01-01

    A short perspective of the Army aeroacoustic research program is presented that emphasizes rotary wing, aerodynamically generated noise. Exciting breakthroughs in experimental techniques and facilities are reviewed which are helping build a detailed understanding of helicopter external noise. Army and joint Army/NASA supported research programs in acoustics which promise to reduce the noise of future helicopters without severe performance penalties are included.

  19. Computational Aero-acoustics As a Tool For Turbo-machinery Noise Reduction

    NASA Technical Reports Server (NTRS)

    Dyson, Rodger W.

    2003-01-01

    This talk will provide an overview of the field of computational aero-acoustics and its use in fan noise prediction. After a brief history of computational fluid dynamics, some of the recent developments in computational aero-acoustics will be explored. Computational issues concerning sound wave production, propagation, and reflection in practical turbo-machinery applications will be discussed including: (a) High order/High Resolution Numerical Techniques. (b) High Resolution Boundary Conditions. [c] MIMD Parallel Computing. [d] Form of Governing Equations Useful for Simulations. In addition, the basic design of our Broadband Analysis Stator Simulator (BASS) code and its application to a 2 D rotor wake-stator interaction will be shown. An example of the noise produced by the wakes from a rotor impinging upon a stator cascade will be shown.

  20. Testing and Performance Verification of a High Bypass Ratio Turbofan Rotor in an Internal Flow Component Test Facility

    NASA Technical Reports Server (NTRS)

    VanZante, Dale E.; Podboy, Gary G.; Miller, Christopher J.; Thorp, Scott A.

    2009-01-01

    A 1/5 scale model rotor representative of a current technology, high bypass ratio, turbofan engine was installed and tested in the W8 single-stage, high-speed, compressor test facility at NASA Glenn Research Center (GRC). The same fan rotor was tested previously in the GRC 9x15 Low Speed Wind Tunnel as a fan module consisting of the rotor and outlet guide vanes mounted in a flight-like nacelle. The W8 test verified that the aerodynamic performance and detailed flow field of the rotor as installed in W8 were representative of the wind tunnel fan module installation. Modifications to W8 were necessary to ensure that this internal flow facility would have a flow field at the test package that is representative of flow conditions in the wind tunnel installation. Inlet flow conditioning was designed and installed in W8 to lower the fan face turbulence intensity to less than 1.0 percent in order to better match the wind tunnel operating environment. Also, inlet bleed was added to thin the casing boundary layer to be more representative of a flight nacelle boundary layer. On the 100 percent speed operating line the fan pressure rise and mass flow rate agreed with the wind tunnel data to within 1 percent. Detailed hot film surveys of the inlet flow, inlet boundary layer and fan exit flow were compared to results from the wind tunnel. The effect of inlet casing boundary layer thickness on fan performance was quantified. Challenges and lessons learned from testing this high flow, low static pressure rise fan in an internal flow facility are discussed.

  1. V/STOL tilt rotor aircraft study. Volume 10: Performance and stability test of A 1-14.622 Froude scaled Boeing Vertol Model 222 tilt rotor aircraft (Phase 1)

    NASA Technical Reports Server (NTRS)

    Mchugh, F. J.; Eason, W.; Alexander, H. R.; Mutter, H.

    1973-01-01

    Wind tunnel test data obtained from a 1/4.622 Froude scale Boeing Model 222 with a full span, two prop, tilt rotor, powered model in the Boeing V/STOL wind tunnel are reported. Data were taken in transition and cruise flight conditions and include performance, stability and control and blade loads information. The effects of the rotors, tail surfaces and airframe on the performance and stability are isolated as are the effects of the airframe on the rotors.

  2. Summary of HEAT 1 Aeroacoustics Installation Effects

    NASA Technical Reports Server (NTRS)

    Smith, Brian E.; Zuniga, Fanny A.; Soderman, Paul T.

    1999-01-01

    A critical part of the NASA High-Speed Research (HSR) program is the demonstration of satisfactory suppression of the jet noise present at low airspeeds. One scheme for reducing jet exhaust noise generated by a future High-Speed Civil Transport (HSCT) is the use of a mixer/ ejector system which would entrain large quantities of ambient air into the exhaust flow from the powerplant in order to cool and slow the jet exhaust before it leaves the tailpipe. Of the variety of factors which can affect the noise suppression characteristics of the mixer/ejector system, the influence of the wing flow field and high-lift devices is not well understood. The effectiveness of the noise suppression device must be evaluated in the presence of the wing/high-lift system before definitive assessments can be made concerning HSCT noise. Of nearly equal importance is the evaluation of the performance of the high-lift system(s) in the presence of realistic propulsion units which feature high ambient flow entrainment rates and jet thrust coefficients. These noise suppressors must provide the required acoustic attenuation while not overly degrading the thrust efficiency of the propulsion system or the lift enhancement of the high-lift devices on the wing. The overall objective of the NASA High-lift Engine Aeroacoustics Technology program is to demonstrate satisfactory interaction between the jet noise suppressor and the high-lift system at airspeeds and angles of attack consistent with takeoff, climb, approach, and landing. In support of this program, an isolated aeroacoustic test of a 13.5%-scale, candidate mixer/ejector nozzle was performed in the Ames' Research Center 40- by 80-Foot Wind Tunnel. The purpose of the test was to measure the baseline aeroacoustic performance characteristics of this nozzle in isolation from the aerodynamic flowfield induced by an HSCT airframe. The test documented the acoustic signature of the nozzles with treated and hardwall ejector surfaces and with

  3. Preliminary Test Results of a Non-Contacting Finger Seal on a Herringbone-Grooved Rotor

    NASA Technical Reports Server (NTRS)

    Proctor, Margaret P.; Delgado, Irebert R.

    2009-01-01

    The baseline non-contacting finger seal is a NASA patented design. The primary difference between it and Gul Aroras design patented by AlliedSignal is that there are no lift pads on the high pressure fingers. The baseline non-contacting finger seal is comprised of a back plate, aft spacer, aft (or low pressure) finger element, forward (or high pressure) finger element, forward spacer, and front plate. The components are held together with 20 flat head screws. A typical seal would have a back plate of approximately the same thickness as the front plate and would be riveted together. The thicker back plate allows use of threaded fasteners so that different finger elements can be tested without having to replace all the individual seal components. The finger elements are essentially washers made of thin sheet stock with multiple curved slots machined around the inner diameter to form the fingers. They are clocked so that the fingers of one cover the slots of the other. The aft finger element fingers have axial extensions or "lift pads" at the seal id that are concentric to the rotor. The fingers act as cantilever beams and flex in response to rotor dynamic motion and radial growth of the rotor due to centrifugal or thermal forces.

  4. Some benchmark problems for computational aeroacoustics

    NASA Astrophysics Data System (ADS)

    Chapman, C. J.

    2004-02-01

    This paper presents analytical results for high-speed leading-edge noise which may be useful for benchmark testing of computational aeroacoustics codes. The source of the noise is a convected gust striking the leading edge of a wing or fan blade at arbitrary subsonic Mach number; the streamwise shape of the gust is top-hat, Gaussian, or sinusoidal, and the cross-stream shape is top-hat, Gaussian, or uniform. Detailed results are given for all nine combinations of shapes; six combinations give three-dimensional sound fields, and three give two-dimensional fields. The gust shapes depend on numerical parameters, such as frequency, rise time, and width, which may be varied arbitrarily in relation to aeroacoustic code parameters, such as time-step, grid size, and artificial viscosity. Hence it is possible to determine values of code parameters suitable for accurate calculation of a given acoustic feature, e.g., the impulsive sound field produced by a gust with sharp edges, or a full three-dimensional acoustic directivity pattern, or a complicated multi-lobed directivity. Another possibility is to check how accurately a code can determine the far acoustic field from nearfield data; a parameter here would be the distance from the leading edge at which the data are taken.

  5. Wind-Tunnel Tests of 10-foot-diameter Autogiro Rotors

    NASA Technical Reports Server (NTRS)

    Wheatley, John B; Bioletti, Carlton

    1937-01-01

    Report presents the results of a series of 10-foot-diameter autogiro rotor models tested in the NACA 20-foot wind tunnel. Four of the models differed only in the airfoil sections of the blades, the sections used being the NACA 0012, 0018, 4412, and 4418. Three additional models employing the NACA 0012 section were tested, in which a varying portion of the blade near the hub was replaced by a streamline tube with a chord of about one-fourth the blade chord.

  6. Halbach Magnetic Rotor Development

    NASA Technical Reports Server (NTRS)

    Gallo, Christopher A.

    2008-01-01

    The NASA John H. Glenn Research Center has a wealth of experience in Halbach array technology through the Fundamental Aeronautics Program. The goals of the program include improving aircraft efficiency, reliability, and safety. The concept of a Halbach magnetically levitated electric aircraft motor will help reduce harmful emissions, reduce the Nation s dependence on fossil fuels, increase efficiency and reliability, reduce maintenance and decrease operating noise levels. Experimental hardware systems were developed in the GRC Engineering Development Division to validate the basic principles described herein and the theoretical work that was performed. A number of Halbach Magnetic rotors have been developed and tested under this program. A separate test hardware setup was developed to characterize each of the rotors. A second hardware setup was developed to test the levitation characteristics of the rotors. Each system focused around a unique Halbach array rotor. Each rotor required original design and fabrication techniques. A 4 in. diameter rotor was developed to test the radial levitation effects for use as a magnetic bearing. To show scalability from the 4 in. rotor, a 1 in. rotor was developed to also test radial levitation effects. The next rotor to be developed was 20 in. in diameter again to show scalability from the 4 in. rotor. An axial rotor was developed to determine the force that could be generated to position the rotor axially while it is rotating. With both radial and axial magnetic bearings, the rotor would be completely suspended magnetically. The purpose of this report is to document the development of a series of Halbach magnetic rotors to be used in testing. The design, fabrication and assembly of the rotors will be discussed as well as the hardware developed to test the rotors.

  7. Aeroacoustic response of coaxial wall-mounted Helmholtz resonators in a low-speed wind tunnel.

    PubMed

    Slaton, William V; Nishikawa, Asami

    2015-01-01

    The aeroacoustic response of coaxial wall-mounted Helmholtz resonators with different neck geometries in a low-speed wind tunnel has been investigated. Experimental test results of this system reveal a strong aeroacoustic response over a Strouhal number range of 0.25 to 0.1 for both increasing and decreasing the flow rate in the wind tunnel. Aeroacoustic response in the low-amplitude range O(10(-3)) < Vac/Vflow < O(10(-1)) has been successfully modeled by describing-function analysis. This analysis, coupled with a turbulent flow velocity distribution model, gives reasonable values for the location in the flow of the undulating stream velocity that drives vortex shedding at the resonator mouth. Having an estimate for the stream velocity that drives the flow-excited resonance is crucial when employing the describing-function analysis to predict aeroacoustic response of resonators.

  8. Evaluation of Rotor Structural and Aerodynamic Loads using Measured Blade Properties

    NASA Technical Reports Server (NTRS)

    Jung, Sung N.; You, Young-Hyun; Lau, Benton H.; Johnson, Wayne; Lim, Joon W.

    2012-01-01

    The structural properties of Higher harmonic Aeroacoustic Rotor Test (HART I) blades have been measured using the original set of blades tested in the wind tunnel in 1994. A comprehensive rotor dynamics analysis is performed to address the effect of the measured blade properties on airloads, blade motions, and structural loads of the rotor. The measurements include bending and torsion stiffness, geometric offsets, and mass and inertia properties of the blade. The measured properties are correlated against the estimated values obtained initially by the manufacturer of the blades. The previously estimated blade properties showed consistently higher stiffnesses, up to 30% for the flap bending in the blade inboard root section. The measured offset between the center of gravity and the elastic axis is larger by about 5% chord length, as compared with the estimated value. The comprehensive rotor dynamics analysis was carried out using the measured blade property set for HART I rotor with and without HHC (Higher Harmonic Control) pitch inputs. A significant improvement on blade motions and structural loads is obtained with the measured blade properties.

  9. Mass Transfer Testing of a 12.5-cm Rotor Centrifugal Contactor

    SciTech Connect

    D. H. Meikrantz; T. G. Garn; J. D. Law; N. R. Mann; T. A. Todd

    2008-09-01

    TRUEX mass transfer tests were performed using a single stage commercially available 12.5 cm centrifugal contactor and stable cerium (Ce) and europium (Eu). Test conditions included throughputs ranging from 2.5 to 15 Lpm and rotor speeds of 1750 and 2250 rpm. Ce and Eu extraction forward distribution coefficients ranged from 13 to 19. The first and second stage strip back distributions were 0.5 to 1.4 and .002 to .004, respectively, throughout the dynamic test conditions studied. Visual carryover of aqueous entrainment in all organic phase samples was estimated at < 0.1 % and organic carryover into all aqueous phase samples was about ten times less. Mass transfer efficiencies of = 98 % for both Ce and Eu in the extraction section were obtained over the entire range of test conditions. The first strip stage mass transfer efficiencies ranged from 75 to 93% trending higher with increasing throughput. Second stage mass transfer was greater than 99% in all cases. Increasing the rotor speed from 1750 to 2250 rpm had no significant effect on efficiency for all throughputs tested.

  10. Takeoff/approach noise for a model counterrotation propeller with a forward-swept upstream rotor

    NASA Technical Reports Server (NTRS)

    Woodward, Richard P.; Hall, David G.; Podboy, Gary G.; Jeracki, Robert J.

    1993-01-01

    A scale model of a counterrotating propeller with forward-swept blades in the forward rotor and aft-swept blades in the aft rotor (designated F39/A31) has been tested in the NASA Lewis 9- by 15-Foot Anechoic Wind Tunnel. This paper presents aeroacoustic results at a takeoff/approach condition of Mach 0.20. Laser Doppler velocimeter results taken in a plane between the two rotors are also included to quantify the interaction flow field. The intention of the forward-swept design is to reduce the magnitude of the forward rotor tip vortex and/or wakes which impinge on the aft rotor, thus lowering the interaction tone levels. A reference model propeller (designated F31/A31), having aft-swept blades in both rotors, was also tested. Aeroelastic performance of the F39/A31 propeller was disappointing. The forward rotor tip region tended to untwist toward higher effective blade angles under load. The forward rotor also exhibited steady state blade flutter at speeds and loadings well below the design condition. The noise results, based on sideline acoustic data, show that the interaction tone levels were up to 8 dB higher with the forward-swept design compared to those for the reference propeller at similar operating conditions, with these tone level differences extending down to lower propeller speeds where flutter did not occur. These acoustic results are for a poorly-performing forward-swept propeller. It is quite possible that a properly-designed forward-swept propeller would exhibit substantial interaction tone level reductions.

  11. Open Rotor Development

    NASA Technical Reports Server (NTRS)

    Van Zante, Dale E.; Rizzi, Stephen A.

    2016-01-01

    The ERA project executed a comprehensive test program for Open Rotor aerodynamic and acoustic performance. System studies used the data to estimate the fuel burn savings and acoustic margin for an aircraft system with open rotor propulsion. The acoustic measurements were used to produce an auralization that compares the legacy blades to the current generation of open rotor designs.

  12. Low-speed wind tunnel test results of the Canard Rotor/Wing concept

    NASA Technical Reports Server (NTRS)

    Bass, Steven M.; Thompson, Thomas L.; Rutherford, John W.; Swanson, Stephen

    1993-01-01

    The Canard Rotor/Wing (CRW), a high-speed rotorcraft concept, was tested at the National Aeronautics and Space Administration (NASA) Ames Research Center's 40- by 80-Foot Wind Tunnel in Mountain View, California. The 1/5-scale model was tested to identify certain low-speed, fixed-wing, aerodynamic characteristics of the configuration and investigate the effectiveness of two empennages, an H-Tail and a T-Tail. The paper addresses the principal test objectives and the results achieved in the wind tunnel test. These are summarized as: i) drag build-up and differences between the H-Tail and T-Tail configuration, ii) longitudinal stability of the H-Tail and T-Tail configurations in the conversion and cruise modes, iii) control derivatives for the canard and elevator in the conversion and cruise modes, iv) aerodynamic characteristics of varying the rotor/wing azimuth position, and v) canard and tail lift/trim capability for conversion conditions.

  13. Computational Aeroacoustic Analysis System Development

    NASA Technical Reports Server (NTRS)

    Hadid, A.; Lin, W.; Ascoli, E.; Barson, S.; Sindir, M.

    2001-01-01

    Many industrial and commercial products operate in a dynamic flow environment and the aerodynamically generated noise has become a very important factor in the design of these products. In light of the importance in characterizing this dynamic environment, Rocketdyne has initiated a multiyear effort to develop an advanced general-purpose Computational Aeroacoustic Analysis System (CAAS) to address these issues. This system will provide a high fidelity predictive capability for aeroacoustic design and analysis. The numerical platform is able to provide high temporal and spatial accuracy that is required for aeroacoustic calculations through the development of a high order spectral element numerical algorithm. The analysis system is integrated with well-established CAE tools, such as a graphical user interface (GUI) through PATRAN, to provide cost-effective access to all of the necessary tools. These include preprocessing (geometry import, grid generation and boundary condition specification), code set up (problem specification, user parameter definition, etc.), and postprocessing. The purpose of the present paper is to assess the feasibility of such a system and to demonstrate the efficiency and accuracy of the numerical algorithm through numerical examples. Computations of vortex shedding noise were carried out in the context of a two-dimensional low Mach number turbulent flow past a square cylinder. The computational aeroacoustic approach that is used in CAAS relies on coupling a base flow solver to the acoustic solver throughout a computational cycle. The unsteady fluid motion, which is responsible for both the generation and propagation of acoustic waves, is calculated using a high order flow solver. The results of the flow field are then passed to the acoustic solver through an interpolator to map the field values into the acoustic grid. The acoustic field, which is governed by the linearized Euler equations, is then calculated using the flow results computed

  14. Further Examination of the Vibratory Loads Reduction Results from the NASA/ARMY/MIT Active Twist Rotor Test

    NASA Technical Reports Server (NTRS)

    Wilbur, Matthew L.; Yeager, William T., Jr.; Sekula, Martin K.

    2002-01-01

    The vibration reduction capabilities of a model rotor system utilizing controlled, strain-induced blade twisting are examined. The model rotor blades, which utilize piezoelectric active fiber composite actuators, were tested in the NASA Langley Transonic Dynamics Tunnel using open-loop control to determine the effect of active-twist on rotor vibratory loads. The results of this testing have been encouraging, and have demonstrated that active-twist rotor designs offer the potential for significant load reductions in future helicopter rotor systems. Active twist control was found to use less than 1% of the power necessary to operate the rotor system and had a pronounced effect on both rotating- and fixed-system loads, offering reductions in individual harmonic loads of up to 100%. A review of the vibration reduction results obtained is presented, which includes a limited set of comparisons with results generated using the second-generation version of the Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics (CAMRAD II) rotorcraft comprehensive analysis.

  15. Validating LES for Jet Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Bridges, James

    2011-01-01

    Engineers charged with making jet aircraft quieter have long dreamed of being able to see exactly how turbulent eddies produce sound and this dream is now coming true with the advent of large eddy simulation (LES). Two obvious challenges remain: validating the LES codes at the resolution required to see the fluid-acoustic coupling, and the interpretation of the massive datasets that result in having dreams come true. This paper primarily addresses the former, the use of advanced experimental techniques such as particle image velocimetry (PIV) and Raman and Rayleigh scattering, to validate the computer codes and procedures used to create LES solutions. It also addresses the latter problem in discussing what are relevant measures critical for aeroacoustics that should be used in validating LES codes. These new diagnostic techniques deliver measurements and flow statistics of increasing sophistication and capability, but what of their accuracy? And what are the measures to be used in validation? This paper argues that the issue of accuracy be addressed by cross-facility and cross-disciplinary examination of modern datasets along with increased reporting of internal quality checks in PIV analysis. Further, it is argued that the appropriate validation metrics for aeroacoustic applications are increasingly complicated statistics that have been shown in aeroacoustic theory to be critical to flow-generated sound.

  16. A multi-frequency fatigue testing method for wind turbine rotor blades

    NASA Astrophysics Data System (ADS)

    Eder, M. A.; Belloni, F.; Tesauro, A.; Hanis, T.

    2017-02-01

    Rotor blades are among the most delicate components of modern wind turbines. Reliability is a crucial aspect, since blades shall ideally remain free of failure under ultra-high cycle loading conditions throughout their designated lifetime of 20-25 years. Full-scale blade tests are the most accurate means to experimentally simulate damage evolution under operating conditions, and are therefore used to demonstrate that a blade type fulfils the reliability requirements to an acceptable degree of confidence. The state-of-the-art testing method for rotor blades in industry is based on resonance excitation where typically a rotating mass excites the blade close to its first natural frequency. During operation the blade response due to external forcing is governed by a weighted combination of its eigenmodes. Current test methodologies which only utilise the lowest eigenfrequency induce a fictitious damage where additional tuning masses are required to recover the desired damage distribution. Even with the commonly adopted amplitude upscaling technique fatigue tests remain a time-consuming and costly endeavour. The application of tuning masses increases the complexity of the problem by lowering the natural frequency of the blade and therefore increasing the testing time. The novel method presented in this paper aims at shortening the duration of the state-of-the-art fatigue testing method by simultaneously exciting the blade with a combination of two or more eigenfrequencies. Taking advantage of the different shapes of the excited eigenmodes, the actual spatial damage distribution can be more realistically simulated in the tests by tuning the excitation force amplitudes rather than adding tuning masses. This implies that in portions of the blade the lowest mode is governing the damage whereas in others higher modes contribute more significantly due to their higher cycle count. A numerical feasibility study based on a publicly available large utility rotor blade is used to

  17. Design, development, and hover testing of a helicopter rotor blade chord extension morphing system

    NASA Astrophysics Data System (ADS)

    Gandhi, Farhan; Hayden, Eric

    2015-03-01

    A rotor blade chord extension system was designed, fabricated and hover tested, using electromechanical and pneumatic actuation. A 1.5 in actuator stroke output in the spanwise direction was converted into chordwise motion of a trailing-edge plate (TEP), via a rigid link. On the hover stand, with a 20 V dc input, the electromechanical actuator was shown to fully extend and retract the plate at rotational speeds up to 385 RPM (which put the system at a centrifugal loading of 209.5 g, or 47.2% of that on a Black Hawk helicopter at 73% span). The configuration was changed to reduce the actuator force requirement for the pneumatic actuator. The rotor test facility allowed a maximum of 105 psi pressure input through the rotary union (significantly lower than the rating of the actuator). At these moderate pressure inputs, full TEP deployment was observed at 315 RPM (140.2 g, or 31.6% of that on a Black Hawk helicopter at 73% span). The model prediction of TEP displacement versus pressure showed good correlation with test results.

  18. Ground and flight test results of a total main rotor isolation system

    NASA Technical Reports Server (NTRS)

    Halwes, Dennis R.

    1987-01-01

    A six degree-of-freedom (DOF) isolation system using six LIVE units has been installed under an Army/NASA contract on a Bell 206LM helicopter. This system has been named the Total Rotor Isolation System, or TRIS. To determine the effectiveness of TRIS in reducing helicopter vibration, a flight verification study was conducted at Bell's Flight Research Center in Arlington, Texas. The flight test data indicate that the 4/rev vibration level at the pilot's seat were suppressed below the 0.04g level throughout the transition envelope. Flight tests indicate over 95% suppression of vibration level from the rotor hub to the pilot's seat. The TRIS installation was designed with a decoupled control system and has shown a significant improvement in aircraft flying qualities, such that it permitted the trimmed aircraft to be flown hands-off for a significant period of time, over 90 seconds. The TRIS flight test program has demonstrated a system that greatly reduces vibration levels of a current-generation helicopter, while significantly improving the flying qualities to a point where stability augmentation is no longer a requirement.

  19. Investigation of helicopter rotor blade/wake interactive impulsive noise

    NASA Technical Reports Server (NTRS)

    Miley, S. J.; Hall, G. F.; Vonlavante, E.

    1987-01-01

    An analysis of the Tip Aerodynamic/Aeroacoustic Test (TAAT) data was performed to identify possible aerodynamic sources of blade/vortex interaction (BVI) impulsive noise. The identification is based on correlation of measured blade pressure time histories with predicted blade/vortex intersections for the flight condition(s) where impulsive noise was detected. Due to the location of the recording microphones, only noise signatures associated with the advancing blade were available, and the analysis was accordingly restricted to the first and second azimuthal quadrants. The results show that the blade tip region is operating transonically in the azimuthal range where previous BVI experiments indicated the impulsive noise to be. No individual blade/vortex encounter is identifiable in the pressure data; however, there is indication of multiple intersections in the roll-up region which could be the origin of the noise. Discrete blade/vortex encounters are indicated in the second quadrant; however, if impulsive noise were produced here, the directivity pattern would be such that it was not recorded by the microphones. It is demonstrated that the TAAT data base is a valuable resource in the investigation of rotor aerodynamic/aeroacoustic behavior.

  20. Open Rotor Tone Shielding Methods for System Noise Assessments Using Multiple Databases

    NASA Technical Reports Server (NTRS)

    Bahr, Christopher J.; Thomas, Russell H.; Lopes, Leonard V.; Burley, Casey L.; Van Zante, Dale E.

    2014-01-01

    Advanced aircraft designs such as the hybrid wing body, in conjunction with open rotor engines, may allow for significant improvements in the environmental impact of aviation. System noise assessments allow for the prediction of the aircraft noise of such designs while they are still in the conceptual phase. Due to significant requirements of computational methods, these predictions still rely on experimental data to account for the interaction of the open rotor tones with the hybrid wing body airframe. Recently, multiple aircraft system noise assessments have been conducted for hybrid wing body designs with open rotor engines. These assessments utilized measured benchmark data from a Propulsion Airframe Aeroacoustic interaction effects test. The measured data demonstrated airframe shielding of open rotor tonal and broadband noise with legacy F7/A7 open rotor blades. Two methods are proposed for improving the use of these data on general open rotor designs in a system noise assessment. The first, direct difference, is a simple octave band subtraction which does not account for tone distribution within the rotor acoustic signal. The second, tone matching, is a higher-fidelity process incorporating additional physical aspects of the problem, where isolated rotor tones are matched by their directivity to determine tone-by-tone shielding. A case study is conducted with the two methods to assess how well each reproduces the measured data and identify the merits of each. Both methods perform similarly for system level results and successfully approach the experimental data for the case study. The tone matching method provides additional tools for assessing the quality of the match to the data set. Additionally, a potential path to improve the tone matching method is provided.

  1. Turbine engine rotor health monitoring evaluation by means of finite element analyses and spin tests data

    NASA Astrophysics Data System (ADS)

    Abdul-Aziz, Ali; Woike, Mark R.; Clem, Michelle; Baaklini, George Y.

    2014-04-01

    Generally, rotating engine components undergo high centrifugal loading environment which subject them to various types of failure initiation mechanisms. Health monitoring of these components is a necessity and is often challenging to implement. This is primarily due to numerous factors including the presence of scattered loading conditions, flaw sizes, component geometry and materials properties, all which hinder the simplicity of applying health monitoring applications. This paper represents a summary work of combined experimental and analytical modeling that included data collection from a spin test experiment of a rotor disk addressing the aforementioned durability issues. It further covers presentation of results obtained from a finite element modeling study to characterize the structural durability of a cracked rotor as it relates to the experimental findings. The experimental data include blade tip clearance, blade tip timing and shaft displacement measurements. The tests were conducted at the NASA Glenn Research Center's Rotordynamics Laboratory, a high precision spin rig. The results are evaluated and examined to determine their significance on the development of a health monitoring system to pre-predict cracks and other anomalies and to assist in initiating a supplemental physics based fault prediction analytical model.

  2. Fan Noise Source Diagnostic Test Computation of Rotor Wake Turbulence Noise

    NASA Technical Reports Server (NTRS)

    Nallasamy, M.; Envia, E.; Thorp, S. A.; Shabbir, A.

    2002-01-01

    An important source mechanism of fan broadband noise is the interaction of rotor wake turbulence with the fan outlet guide vanes. A broadband noise model that utilizes computed rotor flow turbulence from a RANS code is used to predict fan broadband noise spectra. The noise model is employed to examine the broadband noise characteristics of the 22-inch Source Diagnostic Test fan rig for which broadband noise data were obtained in wind tunnel tests at the NASA Glenn Research Center. A 9-case matrix of three outlet guide vane configurations at three representative fan tip speeds are considered. For all cases inlet and exhaust acoustic power spectra are computed and compared with the measured spectra where possible. In general, the acoustic power levels and shape of the predicted spectra are in good agreement with the measured data. The predicted spectra show the experimentally observed trends with fan tip speed, vane count, and vane sweep. The results also demonstrate the validity of using CFD-based turbulence information for fan broadband noise calculations.

  3. Preliminary Test Results of a Non-Contacting Finger Seal on a Herringbone-Grooved Rotor

    NASA Technical Reports Server (NTRS)

    Proctor, Margaret P.; Degado, Irebert R.

    2008-01-01

    Low leakage, non-contacting finger seals have potential to reduce gas turbine engine specific fuel consumption by 2 to 3 percent and to reduce direct operating costs by increasing the time between engine overhauls. A non-contacting finger seal with concentric lift-pads operating adjacent to a test rotor with herringbone grooves was statically tested at 300, 533, and 700 K inlet air temperatures at pressure differentials up to 576 kPa. Leakage flow factors were approximately 70 percent less than state-of-the-art labyrinth seals. Leakage rates are compared to first order predictions. Initial spin tests at 5000 rpm, 300 K inlet air temperature and pressure differentials to 241 kPa produced no measurable wear.

  4. Performance Optimization of a Rotor Alone Nacelle for Acoustic Fan Testing

    NASA Technical Reports Server (NTRS)

    Cunningham, C. C.; Thompson, W. K.; Hughes, C. E.

    2000-01-01

    This paper describes the techniques, equipment, and results from the optimization of a two-axis traverse actuation system used to maintain concentricity between a sting-mounted fan and a wall-mounted nacelle in the 9 x 15 (9 Foot by 15 Foot Test Section) Low Speed Wind Tunnel (LSWT) at the NASA Glenn Research Center (GRC). The Rotor Alone Nacelle (RAN) system, developed at GRC by the Engineering Design and Analysis Division (EDAD) and the Acoustics Branch, used nacelle-mounted lasers and an automated control system to maintain concentricity as thermal and thrust operating loads displace the fan relative to the nacelle. This effort was critical to ensuring rig/facility safety and experimental consistency of the acoustic data from a statorless, externally supported nacelle configuration. Although the tip clearances were originally predicted to be about 0.020 in. at maximum rotor (fan) operating speed, proximity probe measurements showed that the nominal clearance was less than 0.004 in. As a result, the system was optimized through control-loop modifications, active laser cooling, data filtering and averaging, and the development of strict operational procedures. The resultant concentricity error of RAN was reduced to +/- 0.0031 in. in the Y-direction (horizontal) and +0.0035 in./-0.001 3 in. in the Z-direction (vertical), as determined by error analysis and experimental results. Based on the success of this project, the RAN system will be transitioned to other wind tunnel research programs at NASA GRC.

  5. Swept sine testing of rotor-bearing system for damping estimation

    NASA Astrophysics Data System (ADS)

    Chandra, N. Harish; Sekhar, A. S.

    2014-01-01

    Many types of rotating components commonly operate above the first or second critical speed and they are subjected to run-ups and shutdowns frequently. The present study focuses on developing FRF of rotor bearing systems for damping estimation from swept-sine excitation. The principle of active vibration control states that with increase in angular acceleration, the amplitude of vibration due to unbalance will reduce and the FRF envelope will shift towards the right (or higher frequency). The frequency response function (FRF) estimated by tracking filters or Co-Quad analyzers was proved to induce an error into the FRF estimate. Using Fast Fourier Transform (FFT) algorithm and stationary wavelet transform (SWT) decomposition FRF distortion can be reduced. To obtain a theoretical clarity, the shifting of FRF envelope phenomenon is incorporated into conventional FRF expressions and validation is performed with the FRF estimated using the Fourier Transform approach. The half-power bandwidth method is employed to extract damping ratios from the FRF estimates. While deriving half-power points for both types of responses (acceleration and displacement), damping ratio (ζ) is estimated with different approximations like classical definition (neglecting damping ratio of order higher than 2), third order (neglecting damping ratios with order higher than 4) and exact (no assumptions on damping ratio). The use of stationary wavelet transform to denoise the noise corrupted FRF data is explained. Finally, experiments are performed on a test rotor excited with different sweep rates to estimate the damping ratio.

  6. Multi-model Simulation for Optimal Control of Aeroacoustics.

    SciTech Connect

    Collis, Samuel Scott; Chen, Guoquan

    2005-05-01

    Flow-generated noise, especially rotorcraft noise has been a serious concern for bothcommercial and military applications. A particular important noise source for rotor-craft is Blade-Vortex-Interaction (BVI)noise, a high amplitude, impulsive sound thatoften dominates other rotorcraft noise sources. Usually BVI noise is caused by theunsteady flow changes around various rotor blades due to interactions with vorticespreviously shed by the blades. A promising approach for reducing the BVI noise isto use on-blade controls, such as suction/blowing, micro-flaps/jets, and smart struc-tures. Because the design and implementation of such experiments to evaluate suchsystems are very expensive, efficient computational tools coupled with optimal con-trol systems are required to explore the relevant physics and evaluate the feasibilityof using various micro-fluidic devices before committing to hardware.In this thesis the research is to formulate and implement efficient computationaltools for the development and study of optimal control and design strategies for com-plex flow and acoustic systems with emphasis on rotorcraft applications, especiallyBVI noise control problem. The main purpose of aeroacoustic computations is todetermine the sound intensity and directivity far away from the noise source. How-ever, the computational cost of using a high-fidelity flow-physics model across thefull domain is usually prohibitive and itmight also be less accurate because of thenumerical diffusion and other problems. Taking advantage of the multi-physics andmulti-scale structure of this aeroacoustic problem, we develop a multi-model, multi-domain (near-field/far-field) method based on a discontinuous Galerkin discretiza-tion. In this approach the coupling of multi-domains and multi-models is achievedby weakly enforcing continuity of normal fluxes across a coupling surface. For ourinterested aeroacoustics control problem, the adjoint equations that determine thesensitivity of the cost

  7. Fan noise reduction achieved by removing tip flow irregularities behind the rotor - forward arc test configurations

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.; Woodward, R. P.; Mackinnon, M. J.

    1984-01-01

    The noise source caused by the interaction of the rotor tip flow irregularities (vortices and velocity defects) with the downstream stator vanes was studied. Fan flow was removed behind a 0.508 meter (20 in.) diameter model turbofan through an outer wall slot between the rotor and stator. Noise measurements were made with far-field microphones positioned in an arc about the fan inlet and with a pressure transducer in the duct behind the stator. Little tone noise reduction was observed in the forward arc during flow removal; possibly because the rotor-stator interaction noise did not propagate upstream through the rotor. Noise reductions were maded in the duct behind the stator and the largest decrease occurred with the first increment of flow removal. This result indicates that the rotor tip flow irregularity-stator interaction is as important a noise producing mechanism as the normally considered rotor wake-stator interaction.

  8. Low-Speed Wind-Tunnel Test of an Unpowered High-Speed Stoppable Rotor Concept in Fixed-Wing Mode

    NASA Technical Reports Server (NTRS)

    Lance, Michael B.; Sung, Daniel Y.; Stroub, Robert H.

    1991-01-01

    An experimental investigation of the M85, a High Speed Rotor Concept, was conducted at the NASA Langley 14 x 22 foot Subsonic Tunnel, assisted by NASA-Ames. An unpowered 1/5 scale model of the XH-59A helicopter fuselage with a large circular hub fairing, two rotor blades, and a shaft fairing was used as a baseline configuration. The M85 is a rotor wing hybrid aircraft design, and the model was tested with the rotor blade in the fixed wing mode. Assessments were made of the aerodynamic characteristics of various model rotor configurations. Variation in configurations were produced by changing the rotor blade sweep angle and the blade chord length. The most favorable M85 configuration tested included wide chord blades at 0 deg sweep, and it attained a system lift to drag ratio of 8.4.

  9. Laboratory investigation and commercial test for rotors-assembled strand applied in smooth tube

    SciTech Connect

    Li, Feng-Xiang; Ding, Yu-Mei; Guan, Chang-Feng; Xie, Peng-Cheng; Yan, Hua; Yang, Wei-Min

    2008-10-15

    In the commercial test for smooth tube inserted with rotors-assembled strand comparing with non-inserted ones on condensers in electric power plant, using water as working fluid, the single-phase pressure drop and heat transfer were measured. It was found difficult to receive reliable and accurate enough data through commercial test. Meanwhile, the single-phase pressure drop and heat transfer in a rotors-assembled strand inserted tube were measured in laboratory, with the tube side Prandtl numbers varying from 5.67 to 5.80 and the tube side Reynolds numbers varying from 21,300 to 72,200. Before that, a validation experiment based on the same smooth tube was carried out to testify the experimental system and the data reduction method, in which fixed mounts were employed to eliminate entrance effects. The Prandtl numbers varied from 5.64 to 5.76 and the Reynolds numbers varied from 19,000 to 56,000 in the tube. The annular side Reynolds numbers remained nearly constant at the value of around 50,000 for all experiments, with the annular side Prandtl numbers varying from 8.02 to 8.22. The experimental results of smooth tube show that employment of fixed mounts leads to a visible bias of friction factor at relative low Reynolds numbers while it hardly affects the Nusselt numbers. On the other hand, experiment for the tube inserted with rotors-assembled strand show remarkable improvement for heat transfer with the Nusselt number increased by 9.764-11.87% and the overall heat transfer coefficient increased by 7.08-7.49% within the range of Reynolds number from about 21,300 to 55,500. Meanwhile, friction factor increases inevitably by 278.1-353.9% within the same range of Reynolds number. Based on through multivariant linear normal regression method, the Reynolds number and Prandtl number dependencies of the Nusselt number and friction factor were determined to be Nu = 0.0031Re{sup 0.9}Pr{sup 1.0849} and f = 0.993Re{sup -0.22}. (author)

  10. Neutron radiography and other NDE tests of main rotor helicopter blades.

    PubMed

    de Beer, F C; Coetzer, M; Fendeis, D; Da Costa E Silva, A

    2004-10-01

    A few nondestructive examination (NDE) techniques are extensively being used worldwide to investigate aircraft structures for all types of defects. The detection of corrosion and delaminations, which are believed to be the major initiators of defects leading to aircraft structural failures, are addressed by various NDE techniques. In a combined investigation by means of visual inspection, X-ray radiography and shearography on helicopter main rotor blades, neutron radiography (NRad) at SAFARI-1 research reactor operated by Necsa, was performed to introduce this form of NDE testing to the South African aviation industry to be evaluated for applicability. The results of the shearography, visual inspection and NRad techniques are compared in this paper. The main features and advantages of neutron radiography, within the framework of these investigations, will be highlighted.

  11. Investigations of Air-cooled Turbine Rotors for Turbojet Engines II : Mechanical Design, Stress Analysis, and Burst Test of Modified J33 Split-disk Rotor / Richard H. Kemp and Merland L. Moseson

    NASA Technical Reports Server (NTRS)

    Kemp, Richard H; Moseson, Merland L

    1952-01-01

    A full-scale J33 air-cooled split turbine rotor was designed and spin-pit tested to destruction. Stress analysis and spin-pit results indicated that the rotor in a J33 turbojet engine, however, showed that the rear disk of the rotor operated at temperatures substantially higher than the forward disk. An extension of the stress analysis to include the temperature difference between the two disks indicated that engine modifications are required to permit operation of the two disks at more nearly the same temperature level.

  12. Potential acoustic benefits of circulation control rotors

    NASA Technical Reports Server (NTRS)

    Williams, R. M.; Cheeseman, I. C.

    1978-01-01

    The fundamental aeroacoustic mechanisms responsible for noise generation on a rotating blade are theoretically examined. Their contribution to the overall rotor sound pressure level is predicted. Results from a theory for airfoil trailing edge noise are presented. Modifications and extensions to other source theories are described where it is necessary to account for unique aspects of circulation control (CC) aerodynamics. The circulation control rotor (CCR), as embodied on an X-wing vertical takeoff and landing (VTOL) aircraft, is used as an example for computational purposes, although many of the theoretical results presented are generally applicable to other CC applications (such as low speed rotors, propellers, compressors, and fixed wing aircraft). Using the analytical models, it is shown that the utilization CC aerodynamics theoretically makes possible unprecedented advances in rotor noise reduction. For the X-wing VTOL these reductions appear to be feasible without incurring significant attendant performance and weight penalties.

  13. Validating LES for Jet Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Bridges, James; Wernet, Mark P.

    2011-01-01

    Engineers charged with making jet aircraft quieter have long dreamed of being able to see exactly how turbulent eddies produce sound and this dream is now coming true with the advent of large eddy simulation (LES). Two obvious challenges remain: validating the LES codes at the resolution required to see the fluid-acoustic coupling, and the interpretation of the massive datasets that are produced. This paper addresses the former, the use of advanced experimental techniques such as particle image velocimetry (PIV) and Raman and Rayleigh scattering, to validate the computer codes and procedures used to create LES solutions. This paper argues that the issue of accuracy of the experimental measurements be addressed by cross-facility and cross-disciplinary examination of modern datasets along with increased reporting of internal quality checks in PIV analysis. Further, it argues that the appropriate validation metrics for aeroacoustic applications are increasingly complicated statistics that have been shown in aeroacoustic theory to be critical to flow-generated sound, such as two-point space-time velocity correlations. A brief review of data sources available is presented along with examples illustrating cross-facility and internal quality checks required of the data before it should be accepted for validation of LES.

  14. Exhaust System Experiments at NASA's AeroAcoustic Propulsion Lab

    NASA Technical Reports Server (NTRS)

    Bridges, James

    2011-01-01

    This presentation gives an overview of the planned testing in the AeroAcoustic Propulsion Lab (AAPL) in the coming 15 months. It was stressed in the presentation that these are plans that are subject to change due to changes in funding and/or programmatic direction. The first chart shows a simplified schedule of test entries with funding sponsor and dates for each. In subsequent charts are pages devoted to the Objectives and Issues with each test entry, along with a graphic intended to represent the test activity. The chart for each test entry also indicates sponsorship of the activity, and a contact person.!

  15. 14 CFR 29.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... limitations either preclude repeated use of this power or would result in premature engine removals during the... addition, a level of safety equal to that of the main rotors must be provided for— (1) Each component in the rotor drive system whose failure would cause an uncontrolled landing; (2) Each component...

  16. Design and Integration of a Rotor Alone Nacelle for Acoustic Fan Testing

    NASA Technical Reports Server (NTRS)

    Shook, Tony D.; Hughes, Christoper E.; Thompson, William K.; Tavernelli, Paul F.; Cunningham, Cameron C.; Shah, Ashwin

    2001-01-01

    A brief summary of the design, integration and testing of a rotor alone nacelle (RAN) in NASA Glenn's 9'x 15' Low Speed Wind Tunnel (LSWT) is presented. The purpose of the RAN system was to provide an "acoustically clean" flow path within the nacelle to isolate that portion of the total engine system acoustic signature attributed to fan noise. The RAN design accomplished this by removing the stators that provided internal support to the nacelle. In its place, two external struts mounted to a two-axis positioning table located behind the tunnel wall provided the support. Nacelle-mounted lasers and a closed-loop control system provided the input to the table to maintain nacelle to fan concentricity as thermal and thrust loads displaced the strut-mounted fan. This unique design required extensive analysis and verification testing to ensure the safety of the fan model, propulsion simulator drive rig, and facility, along with experimental consistency of acoustic data obtained while using the RAN system. Initial testing was used to optimize the positioning system and resulted in concentricity errors of +/- 0.0031 in. in the horizontal direction and +0.0035/-0.0013 in, in the vertical direction. As a result of successful testing, the RAN system will be transitioned into other acoustic research programs at NASA Glenn Research Center.

  17. Development of flexible rotor balancing criteria

    NASA Technical Reports Server (NTRS)

    Walter, W. W.; Rieger, N. F.

    1979-01-01

    Several studies in which analytical procedures were used to obtain balancing criteria for flexible rotors are described. General response data for a uniform rotor in damped flexible supports were first obtained for plain cylindrical bearings, tilting pad bearings, axial groove bearings, and partial arc bearings. These data formed the basis for the flexible rotor balance criteria presented. A procedure by which a practical rotor in bearings could be reduced to an equivalent uniform rotor was developed and tested. It was found that the equivalent rotor response always exceeded to practical rotor response by more than sixty percent for the cases tested. The equivalent rotor procedure was then tested against six practical rotor configurations for which data was available. It was found that the equivalent rotor method offered a procedure by which balance criteria could be selected for practical flexible rotors, using the charts given for the uniform rotor.

  18. Large-scale Advanced Prop-fan (LAP) static rotor test report

    NASA Technical Reports Server (NTRS)

    Degeorge, Charles L.; Turnberg, Jay E.; Wainauski, Harry S.

    1987-01-01

    Discussed is Static Rotor Testing of the SR-7L Large Scale Advanced Prop-Fan (LAP). The LAP is an advanced 9 foot diameter, 8 bladed propeller designed and built by Hamilton Standard under contract to the NASA Lewis Research Center. The Prop-Fan employs thin swept blades to provide efficient propulsion at flight speeds up to Mach .85. Static Testing was conducted on a 10,000 HP whirl rig at Wright Patterson Air Force Base. The test objectives were to investigate the Prop-Fan static aerodynamic and structural dynamic performance, determine the blade steady state stressers and deflections and to measure steady and unsteady pressures on the SR-7L blade surface. The measured performance of the LAP correlated well with analytical predictions at blade pitch angles below 30 deg. A stall buffet phenomenon was observed at blade pitch angles above 30 deg. This phenomenon manifested itself by elevated blade vibratory stress levels and lower than expected thrust produced and power absorbed by the Prop-Fan for a given speed and blade angle.

  19. Stability Analysis of a High-Speed Seal Test Rotor With Marginal and Extended Squeeze-Film Dampers: Theoretical and Experimental Results

    NASA Technical Reports Server (NTRS)

    Proctor, Margaret P.; Gunter, Edgar J.

    2007-01-01

    A case study of a high-speed seal test rotor shows how rotor dynamic analysis can be used to diagnose the source of high vibrations and evaluate a proposed remedy. Experimental results are compared with the synchronous and non-synchronous whirl response analysis of a double overhung, high-speed seal test rotor with ball bearings supported in 5.84- and 12.7-mm-long, un-centered squeeze-film oil dampers. Test performance with the original damper of length 5.84 mm was marginal. Non-synchronous whirling occurred at the overhung seal test disk and there was a high amplitude synchronous response near the drive spline above 32,000 rpm. Nonlinear synchronous unbalance and time transient whirl studies were conducted on the seal test rotor with the original and extended damper lengths. With the original damper design, the nonlinear synchronous response showed that unbalance could cause damper lockup at 33,000 rpm. Alford cross-coupling forces were also included at the overhung seal test disk for the whirl analysis. Sub-synchronous whirling at the seal test disk was observed in the nonlinear time transient analysis. With the extended damper length of 12.7 mm, the sub-synchronous motion was eliminated and the rotor unbalance response was acceptable to 45,000 rpm with moderate rotor unbalance. Seal test rotor orbits and vibration levels with the extended squeeze film dampers showed smooth operation to 40,444 rpm.

  20. Wind Tunnel Testing of a 120th Scale Large Civil Tilt-Rotor Model in Airplane and Helicopter Modes

    NASA Technical Reports Server (NTRS)

    Theodore, Colin R.; Willink, Gina C.; Russell, Carl R.; Amy, Alexander R.; Pete, Ashley E.

    2014-01-01

    In April 2012 and October 2013, NASA and the U.S. Army jointly conducted a wind tunnel test program examining two notional large tilt rotor designs: NASA's Large Civil Tilt Rotor and the Army's High Efficiency Tilt Rotor. The approximately 6%-scale airframe models (unpowered) were tested without rotors in the U.S. Army 7- by 10-foot wind tunnel at NASA Ames Research Center. Measurements of all six forces and moments acting on the airframe were taken using the wind tunnel scale system. In addition to force and moment measurements, flow visualization using tufts, infrared thermography and oil flow were used to identify flow trajectories, boundary layer transition and areas of flow separation. The purpose of this test was to collect data for the validation of computational fluid dynamics tools, for the development of flight dynamics simulation models, and to validate performance predictions made during conceptual design. This paper focuses on the results for the Large Civil Tilt Rotor model in an airplane mode configuration up to 200 knots of wind tunnel speed. Results are presented with the full airframe model with various wing tip and nacelle configurations, and for a wing-only case also with various wing tip and nacelle configurations. Key results show that the addition of a wing extension outboard of the nacelles produces a significant increase in the lift-to-drag ratio, and interestingly decreases the drag compared to the case where the wing extension is not present. The drag decrease is likely due to complex aerodynamic interactions between the nacelle and wing extension that results in a significant drag benefit.

  1. Turbine Engine Disk Rotor Health Monitoring Assessment Using Spin Tests Data

    NASA Technical Reports Server (NTRS)

    Abdul-Aziz, Ali; Woike, Mark; Baalini, George; Bodis, James R.

    2012-01-01

    Detecting rotating engine component malfunctions and structural anomalies is increasingly becoming a crucial key feature that will help boost safety and lower maintenance cost. However, achievement of such technology, which can be referred to as a health monitoring remains somewhat challenging to implement. This is mostly due to presence of scattered loading conditions, crack sizes, component geometry and material properties that hinders the simplicity of imposing such application. Different approaches are being considered to assist in developing other means of health monitoring or nondestructive techniques to detect hidden flaws and mini cracks before any catastrophic events occur. These methods extend further to assess material discontinuities and other defects that have matured to the level where a failure is very likely. This paper is focused on presenting data obtained from spin test experiments of a turbine engine like rotor disk and their correlation to the development of a structural health monitoring and fault detection system. The data collected includes blade tip clearance, blade tip timing measurements and shaft displacements. The experimental results are collected at rotational speeds up to 10,000 Rpm and tests are conducted at the NASA Glenn Research Center s Rotordynamics Laboratory via a high precision spin system. Additionally, this study offers a closer glance at a selective online evaluation of a rotating disk using advanced capacitive, microwave and eddy current sensor technology.

  2. Turbine engine disk rotor health monitoring assessment using spin tests data

    NASA Astrophysics Data System (ADS)

    Abdul-Aziz, Ali; Woike, Mark; Baaklini, George; Bodis, James R.

    2012-04-01

    Detecting rotating engine component malfunctions and structural anomalies is increasingly becoming a crucial key feature that will help boost safety and lower maintenance cost. However, achievement of such technology, which can be referred to as a health monitoring remains somewhat challenging to implement. This is mostly due to presence of scattered loading conditions, crack sizes, component geometry and material properties that hinders the simplicity of imposing such application. Different approaches are being considered to assist in developing other means of health monitoring or nondestructive techniques to detect hidden flaws and mini cracks before any catastrophic events occur. These methods extend further to assess material discontinuities and other defects that have matured to the level where a failure is very likely. This paper is focused on presenting data obtained from spin test experiments of a turbine engine like rotor disk and their correlation to the development of a structural health monitoring and fault detection system. The data collected includes blade tip clearance, blade tip timing measurements and shaft displacements. The experimental results are collected at rotational speeds up to 10,000 Rpm and tests are conducted at the NASA Glenn Research Center's Rotordynamics Laboratory via a high precision spin system. Additionally, this study offers a closer glance at a selective online evaluation of a rotating disk using advanced capacitive, microwave and eddy current sensor technology.

  3. Design and initial testing of a one-bladed 30-meter-diameter rotor on the NASA/DOE mod-O wind turbine

    NASA Technical Reports Server (NTRS)

    Corrigan, R. D.; Ensworth, C. B. F.

    1986-01-01

    The concept of a one-bladed horizontal-axis wind turbine has been of interest to wind turbine designers for many years. Many designs and economic analyses of one-bladed wind turbines have been undertaken by both United States and European wind energy groups. The analyses indicate significant economic advantages but at the same time, significant dynamic response concerns. In an effort to develop a broad data base on wind turbine design and operations, the NASA Wind Energy Project Office has tested a one-bladed rotor at the NASA/DOE Mod-O Wind Turbine Facility. This is the only known test on an intermediate-sized one-bladed rotor in the United States. The 15.2-meter-radius rotor consists of a tip-controlled blade and a counterweight assembly. A rigorous test series was conducted in the Fall of 1985 to collect data on rotor performance, drive train/generator dynamics, structural dynamics, and structural loads. This report includes background information on one-bladed rotor concepts, and Mod-O one-bladed rotor test configuration, supporting design analysis, the Mod-O one-blade rotor test plan, and preliminary test results.

  4. In-process, non-destructive multimodal dynamic testing of high-speed composite rotors

    NASA Astrophysics Data System (ADS)

    Kuschmierz, Robert; Filippatos, Angelos; Langkamp, Albert; Hufenbach, Werner; Czarske, Jürgern W.; Fischer, Andreas

    2014-03-01

    Fibre reinforced plastic (FRP) rotors are lightweight and offer great perspectives in high-speed applications such as turbo machinery. Currently, novel rotor structures and materials are investigated for the purpose of increasing machine efficiency, lifetime and loading limits. Due to complex rotor structures, high anisotropy and non-linear behavior of FRP under dynamic loads, an in-process measurement system is necessary to monitor and to investigate the evolution of damages under real operation conditions. A non-invasive, optical laser Doppler distance sensor measurement system is applied to determine the biaxial deformation of a bladed FRP rotor with micron uncertainty as well as the tangential blade vibrations at surface speeds above 300 m/s. The laser Doppler distance sensor is applicable under vacuum conditions. Measurements at varying loading conditions are used to determine elastic and plastic deformations. Furthermore they allow to determine hysteresis, fatigue, Eigenfrequency shifts and loading limits. The deformation measurements show a highly anisotropic and nonlinear behavior and offer a deeper understanding of the damage evolution in FRP rotors. The experimental results are used to validate and to calibrate a simulation model of the deformation. The simulation combines finite element analysis and a damage mechanics model. The combination of simulation and measurement system enables the monitoring and prediction of damage evolutions of FRP rotors in process.

  5. Aeroacoustics of Flight Vehicles: Theory and Practice. Volume 1: Noise Sources

    NASA Technical Reports Server (NTRS)

    Hubbard, Harvey H. (Editor)

    1991-01-01

    Methodology recommended to evaluate aeroacoustic related problems is provided, and approaches to their solutions are suggested without extensive tables, nomographs, and derivations. Orientation is toward flight vehicles and emphasis is on underlying physical concepts. Theoretical, experimental, and applied aspects are covered, including the main formulations and comparisons of theory and experiment. The topics covered include: propeller and propfan noise, rotor noise, turbomachinery noise, jet noise classical theory and experiments, noise from turbulent shear flows, jet noise generated by large-scale coherent motion, airframe noise, propulsive lift noise, combustion and core noise, and sonic booms.

  6. Flight Acoustic Testing and Data Acquisition For the Rotor Noise Model (RNM)

    NASA Technical Reports Server (NTRS)

    Conner, David A.; Burley, Casey L.; Smith, Charles D.

    2006-01-01

    Two acoustic flight tests have been conducted on a remote test range at Eglin Air Force Base in the panhandle of Florida. The first was the Acoustics Week flight test conducted in September 2003. The second was the NASA Heavy Lift Rotorcraft Acoustics Flight Test conducted in October-November 2005. Benchmark acoustic databases were obtained for a number of rotorcraft and limited fixed wing vehicles for a variety of flight conditions. The databases are important for validation of acoustic prediction programs such as the Rotorcraft Noise Model (RNM), as well as for the development of low noise flight procedures and for environmental impact assessments. An overview of RNM capabilities and a detailed description of the RNM/ART (Acoustic Repropagation Technique) process are presented. The RNM/ART process is demonstrated using measured acoustic data for the MD600N. The RNM predictions for a level flyover speed sweep show the highest SEL noise levels on the flight track centerline occurred at the slowest vehicle speeds. At these slower speeds, broadband noise content is elevated compared to noise levels obtained at the higher speeds. A descent angle sweep shows that, in general, ground noise levels increased with increasing descent rates. Vehicle orientation in addition to vehicle position was found to significantly affect the RNM/ART creation of source noise semi-spheres for vehicles with highly directional noise characteristics and only mildly affect those with weak acoustic directionality. Based on these findings, modifications are proposed for RNM/ART to more accurately define vehicle and rotor orientation.

  7. Photogrammetric Deflection Measurements for the Tiltrotor Test Rig (TTR) Multi-Component Rotor Balance Calibration

    NASA Technical Reports Server (NTRS)

    Solis, Eduardo; Meyn, Larry

    2016-01-01

    Calibrating the internal, multi-component balance mounted in the Tiltrotor Test Rig (TTR) required photogrammetric measurements to determine the location and orientation of forces applied to the balance. The TTR, with the balance and calibration hardware attached, was mounted in a custom calibration stand. Calibration loads were applied using eleven hydraulic actuators, operating in tension only, that were attached to the forward frame of the calibration stand and the TTR calibration hardware via linkages with in-line load cells. Before the linkages were installed, photogrammetry was used to determine the location of the linkage attachment points on the forward frame and on the TTR calibration hardware. Photogrammetric measurements were used to determine the displacement of the linkage attachment points on the TTR due to deflection of the hardware under applied loads. These measurements represent the first photogrammetric deflection measurements to be made to support 6-component rotor balance calibration. This paper describes the design of the TTR and the calibration hardware, and presents the development, set-up and use of the photogrammetry system, along with some selected measurement results.

  8. Development and Operation of an Automatic Rotor Trim Control System for use During the UH-60 Individual Blade Control Wind Tunnel Test

    NASA Technical Reports Server (NTRS)

    Theodore, Colin R.

    2010-01-01

    A full-scale wind tunnel test to evaluate the effects of Individual Blade Control (IBC) on the performance, vibration, noise and loads of a UH-60A rotor was recently completed in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel [1]. A key component of this wind tunnel test was an automatic rotor trim control system that allowed the rotor trim state to be set more precisely, quickly and repeatably than was possible with the rotor operator setting the trim condition manually. The trim control system was also able to maintain the desired trim condition through changes in IBC actuation both in open- and closed-loop IBC modes, and through long-period transients in wind tunnel flow. This ability of the trim control system to automatically set and maintain a steady rotor trim enabled the effects of different IBC inputs to be compared at common trim conditions and to perform these tests quickly without requiring the rotor operator to re-trim the rotor. The trim control system described in this paper was developed specifically for use during the IBC wind tunnel test

  9. Two-dimensional cascade test of a highly loaded, low-solidity, tandem airfoil turbine rotor blade

    NASA Technical Reports Server (NTRS)

    Kline, J. F.; Stabe, R. G.

    1973-01-01

    A tip region section of a low-solidity tandem airfoil blade for a turbine rotor was tested in a two-dimensional cascade tunnel at solidities of 0.736 and 0.912. Blade surface static pressures and blade exit total and static pressure and flow angle were surveyed. Blade surface velocities, wake shapes, and kinetic energy losses were analyzed and compared with values for 1.852 solidity tandem airfoil blading.

  10. Performance Data from a Wind-Tunnel Test of Two Main-rotor Blade Designs for a Utility-Class Helicopter

    NASA Technical Reports Server (NTRS)

    Singleton, Jeffrey D.; Yeager, William T., Jr.; Wilbur, Matthew L.

    1990-01-01

    An investigation was conducted in the NASA Langley Transonic Dynamics Tunnel to evaluate an advanced main rotor designed for use on a utility class helicopter, specifically the U.S. Army UH-60A Blackhawk. This rotor design incorporated advanced twist, airfoil cross sections, and geometric planform. For evaluation purposes, the current UH-60A main rotor was also tested and is referred to as the baseline blade set. A total of four blade sets were tested. One set of both the baseline and the advanced rotors were dynamically scaled to represent a full scale helicopter rotor blade design. The remaining advanced and baseline blade sets were not dynamically scaled so as to isolate the effects of structural elasticity. The investigation was conducted in hover and at rotor advance ratios ranging from 0.15 to 0.4 at a range of nominal test medium densities from 0.00238 to 0.009 slugs/cu ft. This range of densities, coupled with varying rotor lift and propulsive force, allowed for the simulation of several vehicle gross weight and density altitude combinations. Performance data are presented for all blade sets without analysis; however, cross referencing of data with flight condition may be useful to the analyst for validating aeroelastic theories and design methodologies as well as for evaluating advanced design parameters.

  11. Loads and Performance Data from a Wind-Tunnel Test of Generic Model Helicopter Rotor Blades

    NASA Technical Reports Server (NTRS)

    Yeager, William T., Jr.; Wilbur, Matthew L.

    2005-01-01

    An investigation was conducted in the NASA Langley Transonic Dynamics Tunnel to acquire data for use in assessing the ability of current and future comprehensive analyses to predict helicopter rotating-system and fixed-system vibratory loads. The investigation was conducted with a generic model helicopter rotor system using blades with rectangular planform, no built-in twist, uniform radial distribution of mass and stiffnesses, and a NACA 0012 airfoil section. Rotor performance data, as well as mean and vibratory components of blade bending and torsion moments, fixed-system forces and moments, and pitch link loads were obtained at advance ratios up to 0.35 for various combinations of rotor shaft angle-of-attack and collective pitch. The data are presented without analysis.

  12. An Assessment of NASA Glenn's Aeroacoustic Experimental and Predictive Capabilities for Installed Cooling Fans. Part 1; Aerodynamic Performance

    NASA Technical Reports Server (NTRS)

    VanZante, Dale E.; Koch, L. Danielle; Wernet, Mark P.; Podboy, Gary G.

    2006-01-01

    Driven by the need for low production costs, electronics cooling fans have evolved differently than the bladed components of gas turbine engines which incorporate multiple technologies to enhance performance and durability while reducing noise emissions. Drawing upon NASA Glenn's experience in the measurement and prediction of gas turbine engine aeroacoustic performance, tests have been conducted to determine if these tools and techniques can be extended for application to the aerodynamics and acoustics of electronics cooling fans. An automated fan plenum installed in NASA Glenn's Acoustical Testing Laboratory was used to map the overall aerodynamic and acoustic performance of a spaceflight qualified 80 mm diameter axial cooling fan. In order to more accurately identify noise sources, diagnose performance limiting aerodynamic deficiencies, and validate noise prediction codes, additional aerodynamic measurements were recorded for two operating points: free delivery and a mild stall condition. Non-uniformities in the fan s inlet and exhaust regions captured by Particle Image Velocimetry measurements, and rotor blade wakes characterized by hot wire anemometry measurements provide some assessment of the fan aerodynamic performance. The data can be used to identify fan installation/design changes which could enlarge the stable operating region for the fan and improve its aerodynamic performance and reduce noise emissions.

  13. Lightweight Ceramics for Aeroacoustic Applications

    NASA Technical Reports Server (NTRS)

    Kwan, H. W.; Spamer, G. T.; Yu, J.; Yasukawa, B.

    1997-01-01

    The use of a HTP (High Temperature Performance) ceramic foam for aeroacoustic applications is investigated. HTP ceramic foam is a composition of silica and alumina fibers developed by LMMS. This foam is a lightweight high-temperature fibrous bulk material with small pore size, ultra high porosity, and good strength. It can be used as a broadband noise absorber at both room and high temperature (up to 1800 F). The investigation included an acoustic assessment as well as material development, and environmental and structural evaluations. The results show that the HTP ceramic foam provides good broadband noise absorbing capability and adequate strength when incorporating the HTP ceramic foam system into a honeycomb sandwich structure. On the other hand, the material is sensitive to Skydrol and requires further improvements. Good progress has been made in the impedance model development. A relationship between HTP foam density, flow resistance, and tortuosity will be established in the near future. Additional effort is needed to investigate the coupling effects between face sheet and HTP foam material.

  14. Aeroacoustic Research Techniques: Jets to Autos

    NASA Technical Reports Server (NTRS)

    Soderman, Paul T.

    1999-01-01

    Aeroacoustic research has benefited from the development of advanced techniques for the study of fluid mechanically generated noise New instrumentation; methodologies, information technologies, and facilities have evolved to help researchers investigate the complexities of aircraft and automobile noise. In this paper, research techniques are reviewed with emphasis on the subject closest to the author s experience: aircraft propulsion and airframe noise in simulated flight. A new technology developed for the study of aircraft airframe noise is described as a potential tool for the study of automobile noise. The important role of information technology in aeroacoustic research is discussed.

  15. 14 CFR 29.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... component, except that the control movements need not produce loads or blade flapping motion exceeding the... need not produce loads or blade flapping motion exceeding the maximum loads or motions encountered in... design does not allow stopping the rotors with the engine running, or if no clutch is provided,...

  16. 14 CFR 29.923 - Rotor drive system and control mechanism tests.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... component, except that the control movements need not produce loads or blade flapping motion exceeding the... need not produce loads or blade flapping motion exceeding the maximum loads or motions encountered in... design does not allow stopping the rotors with the engine running, or if no clutch is provided,...

  17. Artificial Icing Test CH-47C Helicopter with Fiberglass Rotor Blades.

    DTIC Science & Technology

    1979-07-01

    fiberglass rotor blades and the intake and exhaust ports for the prototype de-ice system generator . The following general aircraft information is...shut off valve was provided in the tubing run to the generator set fuel inlet. 22 0~0 zgI ogoo j 000 LiJ 2I Photo 3. De-ice System Generator Installation

  18. Optimizing tuning masses for helicopter rotor blade vibration reduction including computed airloads and comparison with test data

    NASA Technical Reports Server (NTRS)

    Pritchard, Jocelyn I.; Adelman, Howard M.; Walsh, Joanne L.; Wilbur, Matthew L.

    1992-01-01

    The development and validation of an optimization procedure to systematically place tuning masses along a rotor blade span to minimize vibratory loads are described. The masses and their corresponding locations are the design variables that are manipulated to reduce the harmonics of hub shear for a four-bladed rotor system without adding a large mass penalty. The procedure incorporates a comprehensive helicopter analysis to calculate the airloads. Predicting changes in airloads due to changes in design variables is an important feature of this research. The procedure was applied to a one-sixth, Mach-scaled rotor blade model to place three masses and then again to place six masses. In both cases the added mass was able to achieve significant reductions in the hub shear. In addition, the procedure was applied to place a single mass of fixed value on a blade model to reduce the hub shear for three flight conditions. The analytical results were compared to experimental data from a wind tunnel test performed in the Langley Transonic Dynamics Tunnel. The correlation of the mass location was good and the trend of the mass location with respect to flight speed was predicted fairly well. However, it was noted that the analysis was not entirely successful at predicting the absolute magnitudes of the fixed system loads.

  19. Flexible rotor dynamics analysis

    NASA Technical Reports Server (NTRS)

    Shen, F. A.

    1973-01-01

    A digital computer program was developed to analyze the general nonaxisymmetric and nonsynchronous transient and steady-state rotor dynamic performance of a bending- and shear-wise flexible rotor-bearing system under various operating conditions. The effects of rotor material mechanical hysteresis, rotor torsion flexibility, transverse effects of rotor axial and torsional loading and the anisotropic, in-phase and out-of-phase bearing stiffness and damping force and moment coefficients were included in the program to broaden its capability. An optimum solution method was found and incorporated in the computer program. Computer simulation of experimental data was made and qualitative agreements observed. The mathematical formulations, computer program verification, test data simulation, and user instruction was presented and discussed.

  20. Acoustic testing of a 1.5 pressure ratio low tip speed fan with a serrated rotor (QEP fan B scale model). [reduction of engine noise

    NASA Technical Reports Server (NTRS)

    Kazin, S. B.; Paas, J. E.; Minzner, W. R.

    1973-01-01

    A scale model of the bypass flow region of a 1.5 pressure ratio, single stage, low tip speed fan was tested with a serrated rotor leading edge to determine its effects on noise generation. The serrated rotor was produced by cutting teeth into the leading edge of the nominal rotor blades. The effects of speed and exhaust nozzle area on the scale models noise characteristics were investigated with both the nominal rotor and serrated rotor. Acoustic results indicate the serrations reduced front quadrant PNL's at takeoff power. In particular, the 200 foot (61.0 m) sideline noise was reduced from 3 to 4 PNdb at 40 deg for nominal and large nozzle operation. However, the rear quadrant maximum sideline PNL's were increased 1.5 to 3 PNdb at approach thust and up to 2 PNdb at takeoff thust with these serrated rotor blades. The configuration with the serrated rotor produced the lowest maximum 200 foot (61.0 m) sideline PNL for any given thust when the large nozzle (116% of design area) was employed.

  1. Rotordynamic Analysis and Feasibility Study of a Disk Spin Test Facility for Rotor Health Monitoring

    NASA Technical Reports Server (NTRS)

    Sawicki, Jerzy T.

    2005-01-01

    Recently, National Aeronautics and Space Administration (NASA) initiated a program to achieve the significant improvement in aviation safety. One of the technical challenges is the design and development of accelerated experiments that mimic critical damage cases encountered in engine components. The Nondestructive Evaluation (NDE) Group at the NASA Glenn Research Center (GRC) is currently addressing the goal concerning propulsion health management and the development of propulsion system specific technologies intended to detect potential failures prior to catastrophe. For this goal the unique disk spin simulation system was assembled at NASA GRC, which allows testing of rotors with the spinning speeds up to 10K RPM, and at the elevated temperature environment reaching 540 C (1000 F). It is anticipated that the facility can be employed for detection of Low Cycle Fatigue disk cracking and further High Cycle Fatigue blade vibration. The controlled crack growth studies at room and elevated temperatures can be conducted on the turbine wheels, and various NDE techniques can be integrated and assessed as in-situ damage monitoring tools. Critical rotating parts in advanced gas turbine engines such as turbine disks frequently operate at high temperature and stress for long periods of time. The integrity of these parts must be proven by non-destructive evaluation (NDE) during various machining steps ranging from forging blank to finished shape, and also during the systematic overhaul inspections. Conventional NDE methods, however, have unacceptable limits. Some of these techniques are time-consuming and inconvenient for service aircraft testing. Almost all of these techniques require that the vicinity of the damage is known in advance. These experimental techniques can provide only local information and no indication of the structural strength at a component and/or system level. The shortcomings of currently available NDE methods lead to the requirement of new damage

  2. Comparison of Theory and Experiment on Aeroacoustic Loads and Deflections

    NASA Astrophysics Data System (ADS)

    Campos, L. M. B. C.; Bourgine, A.; Bonomi, B.

    1999-01-01

    The correlation of acoustic pressure loads induced by a turbulent wake on a nearby structural panel is considered: this problem is relevant to the acoustic fatigue of aircraft, rocket and satellite structures. Both the correlation of acoustic pressure loads and the panel deflections, were measured in an 8-m diameter transonic wind tunnel. Using the measured correlation of acoustic pressures, as an input to a finite-element aeroelastic code, the panel response was reproduced. The latter was also satisfactorily reproduced, using again the aeroelastic code, with input given by a theoretical formula for the correlation of acoustic pressures; the derivation of this formula, and the semi-empirical parameters which appear in it, are included in this paper. The comparison of acoustic responses in aeroacoustic wind tunnels (AWT) and progressive wave tubes (PWT) shows that much work needs to be done to bridge that gap; this is important since the PWT is the standard test means, whereas the AWT is more representative of real flight conditions but also more demanding in resources. Since this may be the first instance of successful modelling of acoustic fatigue, it may be appropriate to list briefly the essential ``positive'' features and associated physical phenomena: (i) a standard aeroelastic structural code can predict acoustic fatigue, provided that the correlation of pressure loads be adequately specified; (ii) the correlation of pressure loads is determined by the interference of acoustic waves, which depends on the exact evaluation of multiple scattering integrals, involving the statistics of random phase shifts; (iii) for the relatively low frequencies (one to a few hundred Hz) of aeroacoustic fatigue, the main cause of random phase effects is scattering by irregular wakes, which are thin on wavelength scale, and appear as partially reflecting rough interfaces. It may also be appropriate to mention some of the ``negative'' features, to which may be attached illusory

  3. Determination of HART I Blade Structural Properties by Laboratory Testing

    NASA Technical Reports Server (NTRS)

    Jung, Sung N.; Lau, Benton H.

    2012-01-01

    The structural properties of higher harmonic Aeroacoustic Rotor Test (HART I) blades were measured using the original set of blades tested in the German-dutch wind tunnel (DNW) in 1994. the measurements include bending and torsion stiffness, geometric offsets, and mass and inertia properties of the blade. the measured properties were compared to the estimated values obtained initially from the blade manufacturer. The previously estimated blade properties showed consistently higher stiffness, up to 30 percent for the flap bending in the blade inboard root section.

  4. An aeroacoustically driven thermoacoustic heat pump

    NASA Astrophysics Data System (ADS)

    Slaton, W. V.; Zeegers, J. C. H.

    2003-04-01

    The mean flow of gas in a pipe past a side branch, closed at the far end, can excite the resonant acoustic modes of the cavity much like blowing across the top of a bottle. This aeroacoustic whistle can excite very high amplitude acoustic waves within the side branch (easily 10% of the mean pressure) at optimal gas flow rates and mean pressures within the main pipe. The aeroacoustic whistle uses no moving parts to convert part of the power in the mean flow into acoustic power. Likewise a thermoacoustic heat pump, utilizing this acoustic power, uses no moving parts to pump heat and establish (or maintain) a temperature difference across a porous medium. This new combination of an aeroacoustic sound source and thermoacoustic heat pump (with suitable thermoelectric elements) is part of an electric power generation feasibility study for natural gas wells. Reliable electrical power generation down-hole to provide electricity for sensors, communications devices or energy storage units is an important research and development goal. Experimental results will be presented that demonstrate the performance of a simple thermoacoustic heat pump when powered by an aeroacoustic sound source. [Work supported by Shell International Exploration and Production B.V.

  5. Mean Flow Boundary Conditions for Computational Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Hixon, R.; Nallasamy, M.; Sawyer, S.; Dyson, R.

    2003-01-01

    In this work, a new type of boundary condition for time-accurate Computational Aeroacoustics solvers is described. This boundary condition is designed to complement the existing nonreflective boundary conditions while ensuring that the correct mean flow conditions are maintained throughout the flow calculation. Results are shown for a loaded 2D cascade, started with various initial conditions.

  6. Aeroacoustics analysis and community noise overview

    NASA Technical Reports Server (NTRS)

    Golub, Robert A.; Soderman, Paul T.

    1992-01-01

    The goals of the High Speed Research Program are focused on three major environmental issues: atmospheric effect, airport community noise, and sonic booms. The issues are basic concerns that require better understanding before further HSRP endeavors can be addresses. This paper discusses airport community noise and aeroacoustic analysis.

  7. Wave rotor demonstrator engine assessment

    NASA Technical Reports Server (NTRS)

    Snyder, Philip H.

    1996-01-01

    The objective of the program was to determine a wave rotor demonstrator engine concept using the Allison 250 series engine. The results of the NASA LERC wave rotor effort were used as a basis for the wave rotor design. A wave rotor topped gas turbine engine was identified which incorporates five basic requirements of a successful demonstrator engine. Predicted performance maps of the wave rotor cycle were used along with maps of existing gas turbine hardware in a design point study. The effects of wave rotor topping on the engine cycle and the subsequent need to rematch compressor and turbine sections in the topped engine were addressed. Comparison of performance of the resulting engine is made on the basis of wave rotor topped engine versus an appropriate baseline engine using common shaft compressor hardware. The topped engine design clearly demonstrates an impressive improvement in shaft horsepower (+11.4%) and SFC (-22%). Off design part power engine performance for the wave rotor topped engine was similarly improved including that at engine idle conditions. Operation of the engine at off design was closely examined with wave rotor operation at less than design burner outlet temperatures and rotor speeds. Challenges identified in the development of a demonstrator engine are discussed. A preliminary design was made of the demonstrator engine including wave rotor to engine transition ducts. Program cost and schedule for a wave rotor demonstrator engine fabrication and test program were developed.

  8. Aeroacoustic sensitivity analysis and optimal aeroacoustic design of turbomachinery blades

    NASA Technical Reports Server (NTRS)

    Hall, Kenneth C.

    1994-01-01

    During the first year of the project, we have developed a theoretical analysis - and wrote a computer code based on this analysis - to compute the sensitivity of unsteady aerodynamic loads acting on airfoils in cascades due to small changes in airfoil geometry. The steady and unsteady flow though a cascade of airfoils is computed using the full potential equation. Once the nominal solutions have been computed, one computes the sensitivity. The analysis takes advantage of the fact that LU decomposition is used to compute the nominal steady and unsteady flow fields. If the LU factors are saved, then the computer time required to compute the sensitivity of both the steady and unsteady flows to changes in airfoil geometry is quite small. The results to date are quite encouraging, and may be summarized as follows: (1) The sensitivity procedure has been validated by comparing the results obtained by 'finite difference' techniques, that is, computing the flow using the nominal flow solver for two slightly different airfoils and differencing the results. The 'analytic' solution computed using the method developed under this grant and the finite difference results are found to be in almost perfect agreement. (2) The present sensitivity analysis is computationally much more efficient than finite difference techniques. We found that using a 129 by 33 node computational grid, the present sensitivity analysis can compute the steady flow sensitivity about ten times more efficiently that the finite difference approach. For the unsteady flow problem, the present sensitivity analysis is about two and one-half times as fast as the finite difference approach. We expect that the relative efficiencies will be even larger for the finer grids which will be used to compute high frequency aeroacoustic solutions. Computational results show that the sensitivity analysis is valid for small to moderate sized design perturbations. (3) We found that the sensitivity analysis provided important

  9. A Subjective Test of Modulated Blade Spacing for Helicopter Main Rotors

    NASA Technical Reports Server (NTRS)

    Sullivan, Brenda M.; Edwards, Bryan D.; Brentner, Kenneth S.; Booth, Earl R., Jr.

    2002-01-01

    Analytically, uneven (modulated) spacing of main rotor blades was found to reduce helicopter noise. A study was performed to see if these reductions transferred to improvements in subjective response. Using a predictive computer code, sounds produced by six main rotor configurations: 4 blades evenly spaced, 5 blades evenly spaced and four configurations with 5 blades with modulated spacing of varying amounts, were predicted. These predictions were converted to audible sounds corresponding to the level flyover, takeoff and approach flight conditions. Subjects who heard the simulations were asked to assess the overflight sounds in terms of noisiness on a scale of 0 to 10. In general the evenly spaced configurations were found less noisy than the modulated spacings, possibly because the uneven spacings produced a perceptible pulsating sound due to the very low fundamental frequency.

  10. Aeroacoustics research in Europe: The CEAS-ASC report on 2009 highlights

    NASA Astrophysics Data System (ADS)

    Casalino, Damiano

    2010-10-01

    The Council of European Aerospace Societies (CEAS) Aeroacoustics Specialists Committee (ASC) supports and promotes the interests of the scientific and industrial aeroacoustics community on an European scale and European aeronautics activities internationally. In this context, "aeroacoustics" encompasses all aerospace acoustics and related areas. Each year the committee highlights some of the research and development projects in Europe. This paper is a report on highlights of aeroacoustics research in Europe in 2009, compiled from information provided to the ASC of the CEAS. In April 2009, the Level-2 project OPENAIR of the Seventh Framework Programme of the European Commission has been launched with the goal of delivering a step change in noise reduction, beyond the successful achievements of the predecessor SILENCE(R). Some contributions submitted to the editor summarizes findings from programmes launched before 2009, while other contributions report on activities supported by national associations and industries. Furthermore, a concise summary of the workshop on "Resolving Uncertainties in Airframe Noise Testing and CAA Code Validation" held in Bucharest is included in this report. Enquiries concerning all contributions should be addressed to the authors who are given at the end of each subsection.

  11. Computational AeroAcoustics for Fan Noise Prediction

    NASA Technical Reports Server (NTRS)

    Envia, Ed; Hixon, Ray; Dyson, Rodger; Huff, Dennis (Technical Monitor)

    2002-01-01

    An overview of the current state-of-the-art in computational aeroacoustics as applied to fan noise prediction at NASA Glenn is presented. Results from recent modeling efforts using three dimensional inviscid formulations in both frequency and time domains are summarized. In particular, the application of a frequency domain method, called LINFLUX, to the computation of rotor-stator interaction tone noise is reviewed and the influence of the background inviscid flow on the acoustic results is analyzed. It has been shown that the noise levels are very sensitive to the gradients of the mean flow near the surface and that the correct computation of these gradients for highly loaded airfoils is especially problematic using an inviscid formulation. The ongoing development of a finite difference time marching code that is based on a sixth order compact scheme is also reviewed. Preliminary results from the nonlinear computation of a gust-airfoil interaction model problem demonstrate the fidelity and accuracy of this approach. Spatial and temporal features of the code as well as its multi-block nature are discussed. Finally, latest results from an ongoing effort in the area of arbitrarily high order methods are reviewed and technical challenges associated with implementing correct high order boundary conditions are discussed and possible strategies for addressing these challenges ore outlined.

  12. Computational test of the infinite order sudden approximation for excitation of linear rigid rotors by collisions with atoms

    NASA Technical Reports Server (NTRS)

    Green, S.

    1978-01-01

    The infinite order sudden approximation for excitation of linear rigid rotors by collisions with atom is tested by comparing integral state-to-state cross sections with accurate close coupling and coupled states results. The systems studied are HCl-Ar, HCl-He, CO-He, HCN-He, CS-H2 and OCS-H2. With the exception of diatomic hydrides (e.g., HCl) which have atypically large rotational constants the method is found to be very accurate to remarkably low collision energies. This approximation should generally be extremely useful for thermal energy collisions.

  13. Application of essentially nonoscillatory methods to aeroacoustic flow problems

    NASA Technical Reports Server (NTRS)

    Atkins, Harold L.

    1995-01-01

    A finite-difference essentially nonoscillatory (ENO) method has been applied to several of the problems prescribed for the workshop sponsored jointly by the Institute for Computer Applications in Science and Engineering and by NASA Langley Research Center entitled 'Benchmark Problems in Computational Aeroacoustics'. The workshop focused on computational challenges specific to aeroacoustics. Among these are long-distance propagation of a short-wavelength disturbance, propagation of small-amplitude disturbances, and nonreflective boundary conditions. The shock capturing-capability inherent to the ENO method effectively eliminates oscillations near shock waves without the need to add and tune dissipation or filter terms. The method-of-lines approach allows the temporal and spatial operators to be chosen separately in accordance with the demands of a particular problem. The ENO method was robust and accurate for all problems in which the propagating wave was resolved with 8 or more points per wavelength. The finite-wave-model boundary condition, a local nonlinear acoustic boundary condition, performed well for the one-dimensional problems. The buffer-domain approach performed well for the two-dimensional test problem. The amplitude of nonphysical reflections were less than 1 percent of the exiting wave's amplitude.

  14. Hover and Wind-Tunnel Testing of Shrouded Rotors for Improved Micro Air Vehicle Design

    DTIC Science & Technology

    2008-01-01

    and Loy, who took me into their family and gave me a home away from home; and my other aunts and uncles — Lisa -Ann, TJ, Bryan, Madeline, Neal and...University XV-11A Marvel. (i) Piasecki 16H-1 Pathfinder . (j) Piasecki 16H-1A Pathfinder II. Figure 1.6: Applications of shrouded rotors: STOL aircraft and...like the Mississippi State University’s XAZ-1 Marvelette and XV-11A Marvel STOL aircraft [77] and the Piasecki 16H-1 Pathfinder and 16H-1A Pathfinder II

  15. Reducing the Effect of Transducer Mount Induced Noise (XMIN) on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

    NASA Technical Reports Server (NTRS)

    Herron, Andrew J.; Reed, Darren K.; Nance, Donald K.

    2015-01-01

    Characterization of flight vehicle unsteady aerodynamics is often studied via large scale wind tunnel testing. Boundary layer noise is measured by miniature pressure transducers installed in a model. Noise levels (2-5 dB ref. 20 µPa) can be induced when transducer is mounted out of flush with model outer surface. This effect must be minimized to accurately determine aerodynamically induced acoustic environments.

  16. Design and preliminary tests of a blade tip air mass injection system for vortex modification and possible noise reduction on a full-scale helicopter rotor

    NASA Technical Reports Server (NTRS)

    Pegg, R. J.; Hosier, R. N.; Balcerak, J. C.; Johnson, H. K.

    1975-01-01

    Full-scale tests were conducted on the Langley helicopter rotor test facility as part of a study to evaluate the effectiveness of a turbulent blade tip air mass injection system in alleviating the impulsive noise (blade slap) caused by blade-vortex interaction. Although blade-slap conditions could not be induced during these tests, qualitative results from flow visualization studies using smoke showed that the differential velocity between the jet velocity and the rotor tip speed was a primary parameter controlling the vortex modification.

  17. Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility. Part 3; Aero-Acoustic Analyses and Experimental Validation

    NASA Technical Reports Server (NTRS)

    Allgood, Daniel C.; Graham, Jason S.; McVay, Greg P.; Langford, Lester L.

    2008-01-01

    A unique assessment of acoustic similarity scaling laws and acoustic analogy methodologies in predicting the far-field acoustic signature from a sub-scale altitude rocket test facility at the NASA Stennis Space Center was performed. A directional, point-source similarity analysis was implemented for predicting the acoustic far-field. In this approach, experimental acoustic data obtained from "similar" rocket engine tests were appropriately scaled using key geometric and dynamic parameters. The accuracy of this engineering-level method is discussed by comparing the predictions with acoustic far-field measurements obtained. In addition, a CFD solver was coupled with a Lilley's acoustic analogy formulation to determine the improvement of using a physics-based methodology over an experimental correlation approach. In the current work, steady-state Reynolds-averaged Navier-Stokes calculations were used to model the internal flow of the rocket engine and altitude diffuser. These internal flow simulations provided the necessary realistic input conditions for external plume simulations. The CFD plume simulations were then used to provide the spatial turbulent noise source distributions in the acoustic analogy calculations. Preliminary findings of these studies will be discussed.

  18. Lifetime Measurements of ^170Hf and a test of the Confined Beta Soft Rotor Model

    NASA Astrophysics Data System (ADS)

    Smith, M. K.; Werner, V.; Heinz, A.; Terry, J. R.; Qian, J.; Winkler, R.; Casperson, R.; Williams, E.; Berant, Z.; L"Uttke, R.; Shoraka, B.; Henning, G.

    2009-10-01

    Significant deviations from rigid rotor model energy level predictions have been known to occur in rare earth nuclei. Recently, it was shown these deviations may be caused by centrifugal stretching effects within the nucleus [1]. New geometrical models have been proposed that account for centrifugal stretching, such as the confined beta soft model (CBS). We present the results from a high precision lifetime experiment performed with the New Yale Plunger Device at WNSL, Yale University. The ground state band of ^170Hf was measured through the J=12^+ level using the Recoil Distance Doppler Shift method. Excited states were populated in the ^124Sn(^50Ti,γ)^170Hf fusion evaporation reaction. Using the lifetimes, the B(E2) values and the quadrupole deformation parameter are determined. Centrifugal stretching is observed as an increased deviation in energy at higher spins in 170Hf. These results are compared to theoretical predictions from the CBS rotor model. Supported by grant DE-FG02-91ER40609.[4pt] [1] Costin et al, Phys.Rev. C 79, 024307 (2009)

  19. A review of ice accretion data from a model rotor icing test and comparison with theory

    NASA Technical Reports Server (NTRS)

    Britton, Randall K.; Bond, Thomas H.

    1991-01-01

    An experiment was conducted by the Helicopter Icing Consortium (HIC) in the NASA Lewis Icing Research Tunnel (IRT) in which a 1/6 scale fuselage model of a UH-60A Black Hawk helicopter with a generic rotor was subjected to a wide range of icing conditions. The HIC consists of members from NASA, Bell Helicopter, Boeing Helicopter, McDonnell Douglas Helicopters, Sikorsky Aircraft, and Texas A&M University. Data was taken in the form of rotor torque, internal force balance measurements, blade strain gage loading, and two dimensional ice shape tracings. A review of the ice shape data is performed with special attention given to repeatability and correctness of trends in terms of radial variation, rotational speed, icing time, temperature, liquid water content, and volumetric median droplet size. Moreover, an indepth comparison between the experimental data and the analysis of NASA's ice accretion code LEWICE is given. Finally, conclusions are shown as to the quality of the ice accretion data and the predictability of the data base as a whole. Recommendations are also given for improving data taking technique as well as potential future work.

  20. A review of ice accretion data from a model rotor icing test and comparison with theory

    NASA Technical Reports Server (NTRS)

    Britton, Randall K.; Bond, Thomas H.

    1991-01-01

    An experiment was conducted by the Helicopter Icing Consortium (HIC) in the NASA Lewis Icing Research Tunnel (IRT) in which a 1/6 scale fuselage model of a UH-60A Black Hawk helicopter with a generic rotor was subjected to a wide range of icing conditions. The HIC consists of members from NASA, Bell Helicopter, Boeing Helicopter, McDonnell Douglas Helicopters, Sikorsky Aircraft, and Texas A&M University. Data was taken in the form of rotor torque, internal force balance measurements, blade strain gage loading, and two dimensional ice shape tracings. A review of the ice shape data is performed with special attention given to repeatability and correctness of trends in terms of radial variation, rotational speed, icing time, temperature, liquid water content, and volumetric median droplet size. Moreover, an indepth comparison between the experimental data and the analysis of NASA's ice accretion code LEWICE is given. Finally, conclusions are drawn as to the quality of the ice accretion data and the predictability of the data base as a whole. Recommendations are also given for improving data taking technique as well as potential future work.

  1. Reducing rotor weight

    SciTech Connect

    Cheney, M.C.

    1997-12-31

    The cost of energy for renewables has gained greater significance in recent years due to the drop in price in some competing energy sources, particularly natural gas. In pursuit of lower manufacturing costs for wind turbine systems, work was conducted to explore an innovative rotor designed to reduce weight and cost over conventional rotor systems. Trade-off studies were conducted to measure the influence of number of blades, stiffness, and manufacturing method on COE. The study showed that increasing number of blades at constant solidity significantly reduced rotor weight and that manufacturing the blades using pultrusion technology produced the lowest cost per pound. Under contracts with the National Renewable Energy Laboratory and the California Energy Commission, a 400 kW (33m diameter) turbine was designed employing this technology. The project included tests of an 80 kW (15.5m diameter) dynamically scaled rotor which demonstrated the viability of the design.

  2. Aeroacoustic Measurements of a Wing-Flap Configuration

    NASA Technical Reports Server (NTRS)

    Meadows, Kristine R.; Brooks, Thomas F.; Humphreys, William M.; Hunter, William H.; Gerhold, Carl H.

    1997-01-01

    Aeroacoustic measurements are being conducted to investigate the mechanisms of sound generation in high-lift wing configurations, and initial results are presented. The model is approximately 6 percent of a full scale configuration, and consists of a main element NACA 63(sub 2) - 215 wing section and a 30 percent chord half-span flap. Flow speeds up to Mach 0.17 are tested at Reynolds number up to approximately 1.7 million. Results are presented for a main element at a 16 degree angle of attack, and flap deflection angles of 29 and 39 degrees. The measurement systems developed for this test include two directional arrays used to localize and characterize the noise sources, and an array of unsteady surface pressure transducers used to characterize wave number spectra and correlate with acoustic measurements. Sound source localization maps show that locally dominant noise sources exist on the flap-side edge. The spectral distribution of the noise sources along the flap-side edge shows a decrease in frequency of the locally dominant noise source with increasing distance downstream of the flap leading edge. Spectra are presented which show general spectral characteristics of Strouhal dependent flow-surface interaction noise. However, the appearance of multiple broadband tonal features at high frequency indicates the presence of aeroacoustic phenomenon following different scaling characteristics. The scaling of the high frequency aeroacoustic phenomenon is found to be different for the two flap deflection angles tested. Unsteady surface pressure measurements in the vicinity of the flap edge show high coherence levels between adjacent sensors on the flap-side edge and on the flap edge upper surface in a region which corresponds closely to where the flap-side edge vortex begins to spill over to the flap upper surface. The frequency ranges where these high levels of coherence occur on the flap surface are consistent with the frequency ranges in which dominant features

  3. Full load shop testing of 18,000-hp gas turbine driven centrifugal compressor for offshore platform service: Evaluation of rotor dynamics performance

    NASA Technical Reports Server (NTRS)

    Kirk, R. G.; Simpson, M.

    1985-01-01

    The results for in-plant full load testing of a 13.4 MW (18000 HP) gas turbine driven centrifugal compressor are presented and compared to analytical predictions of compressor rotor stability. Unique problems from both oil seals and labyrinth gas seals were encountered during the testing. The successful resolution of these problems are summarized.

  4. Flap Edge Aeroacoustic Measurements and Predictions

    NASA Technical Reports Server (NTRS)

    Brooks, Thomas F.; Humphreys, William M., Jr.

    2000-01-01

    An aeroacoustic model test has been conducted to investigate the mechanisms of sound generation on high-lift wing configurations. This paper presents an analysis of flap side-edge noise, which is often the most dominant source. A model of a main element wing section with a half-span flap was tested at low speeds of up to a Mach number of 0.17, corresponding to a wing chord Reynolds number of approximately 1.7 million. Results are presented for flat (or blunt), flanged, and round flap-edge geometries, with and without boundary-layer tripping, deployed at both moderate and high flap angles. The acoustic database is obtained from a Small Aperture Directional Array (SADA) of microphones, which was constructed to electronically steer to different regions of the model and to obtain farfield noise spectra and directivity from these regions. The basic flap-edge aerodynamics is established by static surface pressure data, as well as by Computational Fluid Dynamics (CFD) calculations and simplified edge flow analyses. Distributions of unsteady pressure sensors over the flap allow the noise source regions to be defined and quantified via cross-spectral diagnostics using the SADA output. It is found that shear layer instability and related pressure scatter is the primary noise mechanism. For the flat edge flap, two noise prediction methods based on unsteady-surface-pressure measurements are evaluated and compared to measured noise. One is a new causality spectral approach developed here. The other is a new application of an edge-noise scatter prediction method. The good comparisons for both approaches suggest that much of the physics is captured by the prediction models. Areas of disagreement appear to reveal when the assumed edge noise mechanism does not fully define, the noise production. For the different edge conditions, extensive spectra and directivity are presented. Significantly, for each edge configuration, the spectra for different flow speeds, flap angles, and

  5. Effects of increasing tip velocity on wind turbine rotor design.

    SciTech Connect

    Resor, Brian Ray; Maniaci, David Charles; Berg, Jonathan Charles; Richards, Phillip William

    2014-05-01

    A reduction in cost of energy from wind is anticipated when maximum allowable tip velocity is allowed to increase. Rotor torque decreases as tip velocity increases and rotor size and power rating are held constant. Reduction in rotor torque yields a lighter weight gearbox, a decrease in the turbine cost, and an increase in the capacity for the turbine to deliver cost competitive electricity. The high speed rotor incurs costs attributable to rotor aero-acoustics and system loads. The increased loads of high speed rotors drive the sizing and cost of other components in the system. Rotor, drivetrain, and tower designs at 80 m/s maximum tip velocity and 100 m/s maximum tip velocity are created to quantify these effects. Component costs, annualized energy production, and cost of energy are computed for each design to quantify the change in overall cost of energy resulting from the increase in turbine tip velocity. High fidelity physics based models rather than cost and scaling models are used to perform the work. Results provide a quantitative assessment of anticipated costs and benefits for high speed rotors. Finally, important lessons regarding full system optimization of wind turbines are documented.

  6. Open Rotor: New Option for Jet Engines

    NASA Video Gallery

    NASA's Dale Van Zante describes how the open rotor propulsion system will be tested in a wind tunnel at NASA's Glenn Research Center. Open rotor aircraft engines use high-speed propellers and are c...

  7. Development of a Wind Turbine Test Rig and Rotor for Trailing Edge Flap Investigation: Static Flap Angles Case

    NASA Astrophysics Data System (ADS)

    Abdelrahman, Ahmed; Johnson, David A.

    2014-06-01

    One of the strategies used to improve performance and increase the life-span of wind turbines is active flow control. It involves the modification of the aerodynamic characteristics of a wind turbine blade by means of moveable aerodynamic control surfaces. Trailing edge flaps are relatively small moveable control surfaces placed at the trailing edge of a blade's airfoil that modify the lift of a blade or airfoil section. An instrumented wind turbine test rig and rotor were specifically developed to enable a wide-range of experiments to investigate the potential of trailing edge flaps as an active control technique. A modular blade based on the S833 airfoil was designed to allow accurate instrumentation and customizable settings. The blade is 1.7 meters long, had a constant 178mm chord and a 6° pitch. The modular aerodynamic parts were 3D printed using plastic PC-ABS material. The blade design point was within the range of wind velocities in the available large test facility. The wind facility is a large open jet wind tunnel with a maximum velocity of 11m/s in the test area. The capability of the developed system was demonstrated through an initial study of the effect of stationary trailing edge flaps on blade load and performance. The investigation focused on measuring the changes in flapwise bending moment and power production for different trailing edge flap spanwise locations and deflection angles. The relationship between the load reduction and deflection angle was linear as expected from theory and the highest reduction was caused by the flap furthest from the rotor center. Overall, the experimental setup proved to be effective in measuring small changes in flapwise bending moment within the wind turbine blade and will provide insight when (active) flap control is targeted.

  8. Engine Test Cell Aeroacoustics and Recommendations

    DTIC Science & Technology

    2007-10-01

    SCIPVIX. AIAA J. 23, 505–514. DAVIES, M.G. & OLDFIELD, D.E.S. 1962. Tones from a choked axisymmetric jet. Acustica 12, 257–277. DAVIES, M.G...OLDFIELD, D.E.S. 1962. Tones from a choked axisymmetric jet. PART 2. The self-excited loop and mode of oscillation. Acustica 12, 267–277. HAMMITT, A.G

  9. Effects of aerodynamic interaction between main and tail rotors on helicopter hover performance and noise

    NASA Technical Reports Server (NTRS)

    Menger, R. P.; Wood, T. L.; Brieger, J. T.

    1983-01-01

    A model test was conducted to determine the effects of aerodynamic interaction between main rotor, tail rotor, and vertical fin on helicopter performance and noise in hover out of ground effect. The experimental data were obtained from hover tests performed with a .151 scale Model 222 main rotor, tail rotor and vertical fin. Of primary interest was the effect of location of the tail rotor with respect to the main rotor. Penalties on main rotor power due to interaction with the tail rotor ranged up to 3% depending upon tail rotor location and orientation. Penalties on tail rotor power due to fin blockage alone ranged up to 10% for pusher tail rotors and up to 50% for tractor tail rotors. The main rotor wake had only a second order effect on these tail rotor/fin interactions. Design charts are presented showing the penalties on main rotor power as a function of the relative location of the tail rotor.

  10. Sources, paths, and concepts for reduction of noise in the test section of the NASA Langley 4x7m wind tunnel

    NASA Technical Reports Server (NTRS)

    Hayden, R. E.; Wilby, J. F.

    1984-01-01

    NASA is investigating the feasibility of modifying the 4x7m Wind Tunnel at the Langley Research Center to make it suitable for a variety of aeroacoustic testing applications, most notably model helicopter rotors. The amount of noise reduction required to meet NASA's goal for test section background noise was determined, the predominant sources and paths causing the background noise were quantified, and trade-off studies between schemes to reduce fan noise at the source and those to attenuate the sound generated in the circuit between the sources and the test section were carried out. An extensive data base is also presented on circuit sources and paths.

  11. Effects of the rotor pedalling system on the performance of trained cyclists during incremental and constant-load cycle-ergometer tests.

    PubMed

    Lucía, A; Balmer, J; Davison, R C R; Pérez, M; Santalla, A; Smith, P M

    2004-10-01

    The aim of this study was to determine the effects of Rotor, a new cycle crank configuration that effectively allows the pedals to move independently throughout the duty cycle, on indices of endurance cycling performance in trained cyclists. Ten cyclists (5 Rotor users and 5 non-users; age (mean +/- SD): 22 +/- 5 y; VO(2)max: 69.5 +/- 5.1 mL. kg(-1).min(-1)) volunteered to participate in the study. On four separate days, the subjects performed four cycle-ergometer tests, i.e. two incremental tests and two 20-min tests. An imposed crank rate of 75 rev.min(-1) was used during all tests. The incremental protocol started at 112.5 W, and the power output was increased by 37.5 W every 3 min until volitional exhaustion. The 20-min tests were performed at a fixed power output equivalent to 80 % of the highest power output that the cyclists maintained for a complete 3-min period during incremental tests. Both types of tests were performed with the conventional crank system and the Rotor following a counter-balanced, cross-over design. Gas exchange parameters were measured in all the tests and blood lactate was determined at the end of each 3-min period (incremental tests) and at the end of the 20-min tests. A three factor (pedalling system used during the tests x habitual pedalling system x power output [incremental tests] or time [20-min tests]) ANOVA with repeated measures on power output (incremental tests) or time (20-min tests) was used to analyse several indices of performance, e.g. peak power output, VO(2)max, lactate threshold, onset of blood lactate accumulation, economy, delta, and gross efficiency. No differences (p > 0.05) were found between the Rotor and conventional systems for any of the aforementioned variables. It seems that the theoretical advantage brought about by the Rotor system, i.e. improved contra-lateral cooperation of both legs, would be minimized in trained cyclists. Although field studies are needed to assess the possible implications, in terms

  12. Modeling Aerodynamically Generated Sound of Helicopter Rotors

    NASA Technical Reports Server (NTRS)

    Brentner, Kenneth S.; Farassat, F.

    2002-01-01

    A great deal of progress has been made in the modeling of aerodynamically generated sound of rotors over the past decade. Although the modeling effort has focused on helicopter main rotors, the theory is generally valid for a wide range of rotor configurations. The Ffowcs Williams Hawkings (FW-H) equation has been the foundation for much of the development. The monopole and dipole source terms of the FW-H equation account for the thickness and loading noise, respectively. Bladevortex-interaction noise and broadband noise are important types of loading noise, hence much research has been directed toward the accurate modeling of these noise mechanisms. Both subsonic and supersonic quadrupole noise formulations have been developed for the prediction of high-speed impulsive noise. In an effort to eliminate the need to compute the quadrupole contribution, the FW-H equation has also been utilized on permeable surfaces surrounding all physical noise sources. Comparisons of the Kirchhoff formulation for moving surfaces with the FW-H equation have shown that the Kirchhoff formulation for moving surfaces can give erroneous results for aeroacoustic problems. Finally, significant progress has been made incorporating the rotor noise models into full vehicle noise prediction tools.

  13. Aero-acoustics of Drag Generating Swirling Exhaust Flows

    NASA Technical Reports Server (NTRS)

    Shah, P. N.; Mobed, D.; Spakovszky, Z. S.; Brooks, T. F.; Humphreys, W. M. Jr.

    2007-01-01

    Aircraft on approach in high-drag and high-lift configuration create unsteady flow structures which inherently generate noise. For devices such as flaps, spoilers and the undercarriage there is a strong correlation between overall noise and drag such that, in the quest for quieter aircraft, one challenge is to generate drag at low noise levels. This paper presents a rigorous aero-acoustic assessment of a novel drag concept. The idea is that a swirling exhaust flow can yield a steady, and thus relatively quiet, streamwise vortex which is supported by a radial pressure gradient responsible for pressure drag. Flows with swirl are naturally limited by instabilities such as vortex breakdown. The paper presents a first aero-acoustic assessment of ram pressure driven swirling exhaust flows and their associated instabilities. The technical approach combines an in-depth aerodynamic analysis, plausibility arguments to qualitatively describe the nature of acoustic sources, and detailed, quantitative acoustic measurements using a medium aperture directional microphone array in combination with a previously established Deconvolution Approach for Mapping of Acoustic Sources (DAMAS). A model scale engine nacelle with stationary swirl vanes was designed and tested in the NASA Langley Quiet Flow Facility at a full-scale approach Mach number of 0.17. The analysis shows that the acoustic signature is comprised of quadrupole-type turbulent mixing noise of the swirling core flow and scattering noise from vane boundary layers and turbulent eddies of the burst vortex structure near sharp edges. The exposed edges are the nacelle and pylon trailing edge and the centerbody supporting the vanes. For the highest stable swirl angle setting a nacelle area based drag coefficient of 0.8 was achieved with a full-scale Overall Sound Pressure Level (OASPL) of about 40dBA at the ICAO approach certification point.

  14. Computational approaches to computational aero-acoustics

    NASA Technical Reports Server (NTRS)

    Hardin, Jay C.

    1996-01-01

    The various techniques by which the goal of computational aeroacoustics (the calculation and noise prediction of a fluctuating fluid flow) may be achieved are reviewed. The governing equations for compressible fluid flow are presented. The direct numerical simulation approach is shown to be computationally intensive for high Reynolds number viscous flows. Therefore, other approaches, such as the acoustic analogy, vortex models and various perturbation techniques that aim to break the analysis into a viscous part and an acoustic part are presented. The choice of the approach is shown to be problem dependent.

  15. Classical problems in computational aero-acoustics

    NASA Technical Reports Server (NTRS)

    Hardin, Jay C.

    1996-01-01

    In relation to the expected problems in the development of computational aeroacoustics (CAA), the preliminary applications were to classical problems where the known analytical solutions could be used to validate the numerical results. Such comparisons were used to overcome the numerical problems inherent in these calculations. Comparisons were made between the various numerical approaches to the problems such as direct simulations, acoustic analogies and acoustic/viscous splitting techniques. The aim was to demonstrate the applicability of CAA as a tool in the same class as computational fluid dynamics. The scattering problems that occur are considered and simple sources are discussed.

  16. Theoretical Aeroacoustics: Compiled Mathematical Derivations of Fereidoun 'Feri' Farassat

    NASA Technical Reports Server (NTRS)

    Miller, Steven A. E.

    2016-01-01

    Dr. Fereidoun 'Feri' Farassat was a theoretical aero-acoustician at the National Aeronautics and Space Administration (NASA) Langley Research Center. This book contains technical derivations, notes, and classes that Dr. Farassat produced during his professional career. The layout of the book has been carefully crafted so that foundational ideas through advanced theories, which altered the technical discipline of aeroacoustics, build upon one another. The book can be used to understand the theories of acoustics and learn one contemporary aeroacoustic prediction approach made popular by Dr. Farassat. Most importantly, this book gives the general reader insight into how one of NASA's best aeroacoustics theoreticians thought, constructed, and solved problems throughout his career.

  17. An aeroacoustically driven thermoacoustic heat pump

    NASA Astrophysics Data System (ADS)

    Slaton, W. V.; Zeegers, J. C. H.

    2005-06-01

    The mean flow of gas in a pipe past a cavity can excite the resonant acoustic modes of the cavity-much like blowing across the top of a bottle. The periodic shedding of vortices from the leading edge of the mouth of the cavity feeds energy into the acoustic modes which, in turn, affect the shedding of the next vortex. This so-called aeroacoustic whistle can excite very high amplitude acoustic standing waves within a cavity defined by coaxial side branches closed at their ends. The amplitude of these standing waves can easily be 20% of the ambient pressure at optimal gas flow rates and ambient pressures within the main pipe. A standing wave thermoacoustic heat pump is a device which utilizes the in-phase pressure and displacement oscillations to pump heat across a porous medium thereby establishing, or maintaining, a temperature gradient. Experimental results of a combined system of aeroacoustic sound source and a simple thermoacoustic stack will be presented. .

  18. Open Rotor Noise Prediction at NASA Langley - Capabilities, Research and Development

    NASA Technical Reports Server (NTRS)

    Farassat, Fereidoun

    2010-01-01

    The high fuel prices of recent years have caused the operating cost of the airlines to soar. In an effort to bring down the fuel consumption, the major aircraft engine manufacturers are now taking a fresh look at open rotors for the propulsion of future airliners. Open rotors, also known as propfans or unducted fans, can offer up to 30 per cent improvement in efficiency compared to high bypass engines of 1980 vintage currently in use in most civilian aircraft. NASA Langley researchers have contributed significantly to the development of aeroacoustic technology of open rotors. This report discusses the current noise prediction technology at Langley and reviews the input data requirements, strengths and limitations of each method as well as the associated problems in need of attention by the researchers. We present a brief history of research on the aeroacoustics of rotating blade machinery at Langley Research Center. We then discuss the available noise prediction codes for open rotors developed at NASA Langley and their capabilities. In particular, we present the two useful formulations used for the computation of noise from subsonic and supersonic surfaces. Here we discuss the open rotor noise prediction codes ASSPIN and one based on Ffowcs Williams-Hawkings equation with penetrable data surface (FW - Hpds). The scattering of sound from surfaces near the rotor are calculated using the fast scattering code (FSC) which is also discussed in this report. Plans for further improvements of these codes are given.

  19. Aeroacoustic computation of gust-blade interaction

    NASA Technical Reports Server (NTRS)

    Martin, James E.

    1994-01-01

    To better understand and address the challenges faced in computing the acoustics of flow fields, test problems must be considered. In the present study, the sound radiated by the interaction of a flat plate with an oncoming gust containing a two component, mean velocity is computed. The gust has a uniform mean flow in x with Mach number M(infinity) equal to 0.5. The gust's mean velocity in y is of smaller amplitude and is given by: v = 0.1 sin(pi/8(x/M(sub infinity) - t)). This problem has been posed for an upcoming ICASE/LaRC workshop on benchmark problems in computational aeroacoustics. A plate with a length of 30 units in x is used. The plate is assumed to be infinitesimally thin and is centered at the origin. All variables are made dimensionless using the scales specified. Acoustic quantities are obtained by numerically integrating the linearized Euler equations. Integration is performed on the computational domain -100.0 less than or equal to x less than or equal to 100.0, -100.0 less than or equal to y less than or equal to 100.0, using unit length grid spacing in x and in y. An integration scheme is sought which will provide accurate solution to the small quantities of interest at a minimal computational expense. Results indicate that with the given discretization a scheme of minimal fourth order accuracy might be adequate to approximate the waves within the given flow. Thus, a variation of the MacCormack scheme with fourth order accuracy in space and second order accuracy in time was chosen. A scheme with sixth order accuracy in space has also been implemented and results compared with those of the fourth order accurate scheme. To ensure no mass flux, zero normal velocity is assigned at the plate. This condition will induce a discontinuity in the pressure across the plate location. Values for the perturbation pressure p' along the surface of the plate are obtained using a one-sided, third order Taylor expansion, such that p'(sub y) = O. In accordance with

  20. Correlating CFD Simulation with Wind Tunnel Test for the Full-Scale UH-60A Airloads Rotor

    NASA Technical Reports Server (NTRS)

    Romandr, Ethan; Norman, Thomas R.; Chang, I-Chung

    2011-01-01

    Data from the recent UH-60A Airloads Test in the National Full-Scale Aerodynamics Complex 40- by 80- Foot Wind Tunnel at NASA Ames Research Center are presented and compared to predictions computed by a loosely coupled Computational Fluid Dynamics (CFD)/Comprehensive analysis. Primary calculations model the rotor in free-air, but initial calculations are presented including a model of the tunnel test section. The conditions studied include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed into deep stall. Predictions show reasonable agreement with measurement for integrated performance indicators such as power and propulsive but occasionally deviate significantly. Detailed analysis of sectional airloads reveals good correlation in overall trends for normal force and pitching moment but pitching moment mean often differs. Chord force is frequently plagued by mean shifts and an overprediction of drag on the advancing side. Locations of significant aerodynamic phenomena are predicted accurately although the magnitude of individual events is often missed.

  1. Investigation of rotor blade element airloads for a teetering rotor in the blade stall regime

    NASA Technical Reports Server (NTRS)

    Dadone, L. U.; Fukushima, T.

    1974-01-01

    A model of a teetering rotor was tested in a low speed wind tunnel. Blade element airloads measured on an articulated model rotor were compared with the teetering rotor and showed that the teetering rotor is subjected to less extensive flow separation. Retreating blade stall was studied. Results show that stall, under the influence of unsteady aerodynamic effects, consists of four separate stall events, each associated with a vortex shed from the leading edge and sweeping over the upper surface of the rotor blade. Current rotor performance prediction methodology was evaluated through computer simulation.

  2. Molecular Rotors

    DTIC Science & Technology

    2006-10-31

    Zack Murray No Faith Jordan No Charles Hagerdorn No Elizabeth Baker, No Mathew Myers No Leah A. Edelman No William T. Gathright No Bill Bockrath No...ligand could impact the construction of rotor arrays. Consequently, VASP calculations were performed in collaboration with Dr. Jeffrey Reimers of

  3. Separators for flywheel rotors

    DOEpatents

    Bender, D.A.; Kuklo, T.C.

    1998-07-07

    A separator forms a connection between the rotors of a concentric rotor assembly. This separator allows for the relatively free expansion of outer rotors away from inner rotors while providing a connection between the rotors that is strong enough to prevent disassembly. The rotor assembly includes at least two rotors referred to as inner and outer flywheel rings or rotors. This combination of inner flywheel ring, separator, and outer flywheel ring may be nested to include an arbitrary number of concentric rings. The separator may be a segmented or continuous ring that abuts the ends of the inner rotor and the inner bore of the outer rotor. It is supported against centrifugal loads by the outer rotor and is affixed to the outer rotor. The separator is allowed to slide with respect to the inner rotor. It is made of a material that has a modulus of elasticity that is lower than that of the rotors. 10 figs.

  4. Separators for flywheel rotors

    DOEpatents

    Bender, Donald A.; Kuklo, Thomas C.

    1998-01-01

    A separator forms a connection between the rotors of a concentric rotor assembly. This separator allows for the relatively free expansion of outer rotors away from inner rotors while providing a connection between the rotors that is strong enough to prevent disassembly. The rotor assembly includes at least two rotors referred to as inner and outer flywheel rings or rotors. This combination of inner flywheel ring, separator, and outer flywheel ring may be nested to include an arbitrary number of concentric rings. The separator may be a segmented or continuous ring that abuts the ends of the inner rotor and the inner bore of the outer rotor. It is supported against centrifugal loads by the outer rotor and is affixed to the outer rotor. The separator is allowed to slide with respect to the inner rotor. It is made of a material that has a modulus of elasticity that is lower than that of the rotors.

  5. Full scale wind tunnel investigation of a bearingless main helicopter rotor. [Ames 40 by 80 foot wind tunnel test using the BO-105 helicopter

    NASA Technical Reports Server (NTRS)

    1980-01-01

    A stability test program was conducted to determine the effects of airspeed, collective pitch, rotor speed and shaft angle on stability and loads at speeds beyond that attained in the BMR/BO-105 flight test program. Loads and performance data were gathered at forward speeds up to 165 knots. The effect of cyclic pitch perturbations on rotor response was investigated at simulated level flight conditions. Two configuration variations were tested for their effect on stability. One variable was the control system stiffness. An axially softer pitch link was installed in place of the standard BO-105 pitch link. The second variation was the addition of elastomeric damper strips to increase the structural damping. The BMR was stable at all conditions tested. At fixed collective pitch, shaft angle and rotor speed, damping generally increased between hover and 60 knots, remained relatively constant from 60 to 90 knots, then decreased above 90 knots. Analytical predictions are in good agreement with test data up to 90 knots, but the trend of decreasing damping above 90 knots is contrary to the theory.

  6. Highlights of Aeroacoustics Research in the U.S. 1998

    NASA Technical Reports Server (NTRS)

    Raman, Ganesh; McLaughlin, Dennis K.

    1999-01-01

    Highlights of aeroacoustics research in the United States of America during 1998 are reported in a summary compiled from information provided by members of the Aeroacoustics Technical Committee of the American Institute of Aeronautics and Astronautics (AIAA) and other leading research groups in industry, national laboratories, and academia. The past few years have seen significant progress in aeroacoustics. Research has steadily progressed toward enhanced safety, noise benefits, and lower costs. Since industrial progress is generally not published in the archival literature, it is particularly important to highlight these accomplishments. This year we chose to report on five topics of great interest to the aerospace industry including a synopsis of fundamental research at universities and national laboratories. The topics chosen are: (1) Advanced Subsonic Technology (AST), (2) High Speed Research (HSR), (3) Rotorcraft, (4) Weapons bay aeroacoustics control and (5) Academic research including Computational AeroAcoustics (CAA). Although the information presented in this review is not all encompassing we hope that the topics covered will provide some insights into aeroacoustics activity in the U.S.

  7. Full-Scale Wind Tunnel Test of the UH-60A Airloads Rotor

    DTIC Science & Technology

    2011-05-01

    acquisition and reduction systems. In addition, the data validation process, the test objectives and approach, and some sample results are presented. The...the data validation process, the test objectives and approach, and some sample results are presented. The test has produced unique data not available...SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT Same as Report (SAR) 18. NUMBER OF PAGES 26 19a. NAME OF RESPONSIBLE PERSON a. REPORT

  8. Aeroacoustic simulation for phonation modeling

    NASA Astrophysics Data System (ADS)

    Irwin, Jeffrey; Hanford, Amanda; Craven, Brent; Krane, Michael

    2011-11-01

    The phonation process occurs as air expelled from the lungs creates a pressure drop and a subsequent air flow across the larynx. The fluid-structure interaction between the turbulent air flow and oscillating vocal folds, combined with additional resonance in the oral and nasal cavities, creates much of what we hear in the human voice. As many voice-related disorders can be traced to irregular vocal tract shape or motion, it is important to understand in detail the physics involved in the phonation process. To numerically compute the physics of phonation, a solver must be able to accurately model acoustic airflow through a moving domain. The open-source CFD package OpenFOAM is currently being used to evaluate existing solvers against simple acoustic test cases, including an open-ended resonator and an expansion chamber, both of which utilize boundary conditions simulating acoustic sources as well as anechoic termination. Results of these test cases will be presented and compared with theory, and the future development of a three-dimensional vocal tract model and custom-mode acoustic solver will be discussed. Acknowledge support of NIH grant 5R01DC005642 and ARL E&F program.

  9. Comparison of analytical and flight test identified aerodynamic derivatives for a tandem-rotor transport helicopter

    NASA Technical Reports Server (NTRS)

    Hodge, W. F.

    1980-01-01

    Flight tests for verifying an analytical aerodynamic derivative model of a CH-47 helicopter were conducted for low cruise speeds and transition to hover portions of curved, decelerating landing approach trajectories. All testing was performed on a closed loop basis with the stability augmentation system of the helicopter operating, and response data were obtained using both manual and computer generated input maneuvers. The results indicate some differences between the measured response time histories and those predicted by both analytical and flight test identified derivatives. With some exceptions the discrepancies are not severe, and the overall agreement between the measured and computed time histories is reasonably good. No adverse effects attributable to closed loop testing were noted, and the use of computer generated inputs proved to be superior to manual ones.

  10. Multimodel methods for optimal control of aeroacoustics.

    SciTech Connect

    Chen, Guoquan; Collis, Samuel Scott

    2005-01-01

    A new multidomain/multiphysics computational framework for optimal control of aeroacoustic noise has been developed based on a near-field compressible Navier-Stokes solver coupled with a far-field linearized Euler solver both based on a discontinuous Galerkin formulation. In this approach, the coupling of near- and far-field domains is achieved by weakly enforcing continuity of normal fluxes across a coupling surface that encloses all nonlinearities and noise sources. For optimal control, gradient information is obtained by the solution of an appropriate adjoint problem that involves the propagation of adjoint information from the far-field to the near-field. This computational framework has been successfully applied to study optimal boundary-control of blade-vortex interaction, which is a significant noise source for helicopters on approach to landing. In the model-problem presented here, the noise propagated toward the ground is reduced by 12dB.

  11. Analysis and Testing of a Composite Fuselage Shield for Open Rotor Engine Blade-Out Protection

    NASA Technical Reports Server (NTRS)

    Pereira, J. Michael; Emmerling, William; Seng, Silvia; Frankenberger, Charles; Ruggeri, Charles R.; Revilock, Duane M.; Carney, Kelly S.

    2015-01-01

    The Federal Aviation Administration is working with the European Aviation Safety Agency to determine the certification base for proposed new engines that would not have a containment structure on large commercial aircraft. Equivalent safety to the current fleet is desired by the regulators, which means that loss of a single fan blade will not cause hazard to the Aircraft. The NASA Glenn Research Center and The Naval Air Warfare Center (NAWC), China Lake, collaborated with the FAA Aircraft Catastrophic Failure Prevention Program to design and test lightweight composite shields for protection of the aircraft passengers and critical systems from a released blade that could impact the fuselage. In the test, two composite blades were pyrotechnically released from a running engine, each impacting a composite shield with a different thickness. The thinner shield was penetrated by the blade and the thicker shield prevented penetration. This was consistent with pre-test predictions. This paper documents the live fire test from the full scale rig at NAWC China Lake and describes the damage to the shields as well as instrumentation results.

  12. Analysis and Testing of a Composite Fuselage Shield for Open Rotor Engine Blade-Out Protection

    NASA Technical Reports Server (NTRS)

    Pereira, J. Michael; Emmerling, William; Seng, Silvia; Frankenberger, Charles; Ruggeri, Charles R.; Revilock, Duane M.; Carney, Kelly S.

    2016-01-01

    The Federal Aviation Administration is working with the European Aviation Safety Agency to determine the certification base for proposed new engines that would not have a containment structure on large commercial aircraft. Equivalent safety to the current fleet is desired by the regulators, which means that loss of a single fan blade will not cause hazard to the Aircraft. The NASA Glenn Research Center and The Naval Air Warfare Center (NAWC), China Lake, collaborated with the FAA Aircraft Catastrophic Failure Prevention Program to design and test lightweight composite shields for protection of the aircraft passengers and critical systems from a released blade that could impact the fuselage. LS-DYNA® was used to predict the thickness of the composite shield required to prevent blade penetration. In the test, two composite blades were pyrotechnically released from a running engine, each impacting a composite shield with a different thickness. The thinner shield was penetrated by the blade and the thicker shield prevented penetration. This was consistent with pre-test LS-DYNA predictions. This paper documents the analysis conducted to predict the required thickness of a composite shield, the live fire test from the full scale rig at NAWC China Lake and describes the damage to the shields as well as instrumentation results.

  13. 14 CFR 33.94 - Blade containment and rotor unbalance tests.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.94... blade must fail. (2) Failure of the most critical turbine blade while operating at maximum permissible r... discs, at least 80 percent of the blade must fail. The most critical turbine blade must be determined...

  14. Aeroacoustic Evaluation of Flap and Landing Gear Noise Reduction Concepts

    NASA Technical Reports Server (NTRS)

    Khorrami, Mehdi R.; Humphreys, William M., Jr.; Lockard, David P.; Ravetta, Patricio A.

    2014-01-01

    Aeroacoustic measurements for a semi-span, 18% scale, high-fidelity Gulfstream aircraft model are presented. The model was used as a test bed to conduct detailed studies of flap and main landing gear noise sources and to determine the effectiveness of numerous noise mitigation concepts. Using a traversing microphone array in the flyover direction, an extensive set of acoustic data was obtained in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel with the facility in the acoustically treated open-wall (jet) mode. Most of the information was acquired with the model in a landing configuration with the flap deflected 39 deg and the main landing gear alternately installed and removed. Data were obtained at Mach numbers of 0.16, 0.20, and 0.24 over directivity angles between 56 deg and 116 deg, with 90 deg representing the overhead direction. Measured acoustic spectra showed that several of the tested flap noise reduction concepts decrease the sound pressure levels by 2 - 4 dB over the entire frequency range at all directivity angles. Slightly lower levels of noise reduction from the main landing gear were obtained through the simultaneous application of various gear devices. Measured aerodynamic forces indicated that the tested gear/flap noise abatement technologies have a negligible impact on the aerodynamic performance of the aircraft model.

  15. Exploratory Engine Test of Transpiration-cooled Turbine-rotor Blade with Wire-cloth Shell

    NASA Technical Reports Server (NTRS)

    Donoughe, Patrick L; Diaguila, Anthony J

    1954-01-01

    Engine tests were made on a transpiration-cooled blade that was fabricated from an internal load-carrying member with an external surface of wire cloth. After operation in the engine, some damage was noted at the tip region of the trailing edge of the blades. On other sections of the blade, the wire cloth did not appear greatly overheated, and it appeared that satisfactory chordwise temperature distribution was provided by orifices in the blade base.

  16. 76 FR 42020 - Airworthiness Standards; Rotor Overspeed Requirements

    Federal Register 2010, 2011, 2012, 2013, 2014

    2011-07-18

    ... ``extremely remote'' in Sec. 33.27(c); Exclusions of shaft sections from overspeed tests; Material properties... comment. Material Properties of Test Rotors Section 33.27(a)(1) proposed that test rotors used to... validated to prior overspeed test results of a similar rotor. The tool must be validated for each...

  17. An examination of the aerodynamic moment on rotor blade tips using flight test data and analysis

    NASA Technical Reports Server (NTRS)

    Maier, Thomas H.; Bousman, William G.

    1993-01-01

    The analysis CAMRAD/JA is used to model two aircraft, a Puma with a swept-tip blade and a UH-60A Black Hawk. The accuracy of the analysis in predicting the torsion loads is assessed by comparing the predicted loads with measurements from flight tests. The influence of assumptions in the analytical model is examined by varying model parameters and comparing the predicted results to baseline values for the torsion loads. Flight test data from a research Puma are used to identify the source of torsion loads. These data indicate that the aerodynamic section moment in the region of the blade tip dominates torsion loading in high-speed flight. Both the aerodynamic section moment at the blade tip and the pitch-link loads are characterized by large positive (nose-up) moments in the first quadrant with rapid reversal of load so that the moment is negative in the second quadrant. Both the character and magnitude of this loading are missed by the CAMRAD/JA analysis.

  18. An immersed boundary method for aeroacoustic flow using a high-order finite difference method

    NASA Astrophysics Data System (ADS)

    Olson, Britton

    2016-11-01

    An immersed boundary method that achieves second order accuracy in space on acoustic reflection problems is introduced and tested on a number of aero-acoustic related problems. The method follows a continuous forcing approach and uses existing solver operators to smoothly extend the flow solution though the immersed boundary. Both no-slip and free-slip boundary conditions are demonstrated on complex geometries using a high-order finite difference code on a Cartesian grid. High Mach number test problems are also shown, demonstrating the method's robustness in the presence of shock waves.

  19. Wind tunnel testing of 5-bladed H-rotor wind turbine with the integration of the omni-direction-guide-vane

    NASA Astrophysics Data System (ADS)

    Fazlizan, A.; Chong, W. T.; Omar, W. Z. W.; Mansor, S.; Zain, Z. M.; Pan, K. C.; Oon, C. S.

    2012-06-01

    A novel omni-direction-guide-vane (ODGV) that surrounds a vertical axis wind turbine (VAWT) is designed to improve the wind turbine performance by increasing the oncoming wind speed and guiding the wind-stream through optimum flow angles before impinging onto the turbine blades. Wind tunnel testing was performed to measure the performance of a 5-bladed H-rotor wind turbine with Wortmann FX63-137 airfoil blades, with and without the integration of the ODGV. The test was conducted using a scaled model turbine which was constructed to simulate the VAWT enclosed by the ODGV on a building. The diameter and height of the ODGV are 2 times larger than the VAWT's. Torque, rotational speed and power measurements were performed by using torque transducer with hysteresis brake applied to the rotor shaft. The VAWT shows an improvement on its self-starting behavior where the cut-in speed reduced to 4 m/s with the ODGV (7.35 m/s without the ODGV). Since the VAWT is able to self-start at lower wind speed, the working hour of the wind turbine would increase. At the wind speed of 6 m/s and free-running condition (only rotor inertia and bearing friction were applied), the ODGV helps to increase the rotor RPM by 182%. At the same wind speed (6 m/s), the ODGV helps to increase the power output by 3.48 times at peak torque. With this innovative design, the size of VAWT can be reduced for a given power output and should generate interest in the market, even for regions with weaker winds.

  20. Preliminary design of mesoscale turbocompressor and rotordynamics tests of rotor bearing system

    NASA Astrophysics Data System (ADS)

    Hossain, Md Saddam

    2011-12-01

    A mesoscale turbocompressor spinning above 500,000 RPM is evolutionary technology for micro turbochargers, turbo blowers, turbo compressors, micro-gas turbines, auxiliary power units, etc for automotive, aerospace, and fuel cell industries. Objectives of this work are: (1) to evaluate different air foil bearings designed for the intended applications, and (2) to design & perform CFD analysis of a micro-compressor. CFD analysis of shrouded 3-D micro compressor was conducted using Ansys Bladegen as blade generation tool, ICEM CFD as mesh generation tool, and CFX as main solver for different design and off design cases and also for different number of blades. Comprehensive experimental facilities for testing the turbocompressor system have been also designed and proposed for future work.

  1. Evaluation of Demo 1C composite flywheel rotor burst test and containment design

    SciTech Connect

    Kass, M.D.; McKeever, J.W.; Akerman, M.A.; Goranson, P.L.; Litherland, P.S.; O`Kain, D.U.

    1998-07-01

    Laboratory-Directed funds were provided in FY 1995 for research to develop flywheel containment specifications and to consider concepts that could satisfy these specifications and produce a prototype small, lightweight, inexpensive, mobile flywheel containment. Research activities have included an analytical and pictorial review of the Demo 1C flywheel failure test, which provided significant insight about radial and axial failure modes; calculations of the thickness of ultra-conservative pressure vessel containment; entertainment of advanced containment concepts using lightweight materials and armor literature; consideration of fabrication assembly procedures; and participation in a Flywheel Energy Storage Workshop during which additional flywheel failure experiences were discussed. Based on these activities, calculations, and results, a list of conclusions concerning flywheel containment and its relation to the flywheel are presented followed by recommendations for further research.

  2. NASA Open Rotor Noise Research

    NASA Technical Reports Server (NTRS)

    Envia, Ed

    2010-01-01

    Owing to their inherent fuel burn efficiency advantage compared with the current generation high bypass ratio turbofan engines, there is resurgent interest in developing open rotor propulsion systems for powering the next generation commercial aircraft. However, to make open rotor systems truly competitive, they must be made to be acoustically acceptable too. To address this challenge, NASA in collaboration with industry is exploring the design space for low-noise open rotor propulsion systems. The focus is on the system level assessment of the open rotors compared with other candidate concepts like the ultra high bypass ratio cycle engines. To that end there is an extensive research effort at NASA focused on component testing and diagnostics of the open rotor acoustic performance as well as assessment and improvement of open rotor noise prediction tools. In this presentation and overview of the current NASA research on open rotor noise will be provided. Two NASA projects, the Environmentally Responsible Aviation Project and the Subsonic Fixed Wing Project, have been funding this research effort.

  3. Rotor fatigue monitoring data acquisition system

    NASA Technical Reports Server (NTRS)

    Smith, Scott M.

    1993-01-01

    The 40 by 80 Foot Wind Tunnel of the National Full Scale Aerodynamics Complex (NFAC) had a requirement to monitor rotor fatigue during a test. This test subjected various rotor components to stress levels higher than their structural fatigue limits. A data acquisition system was developed to monitor the cumulative fatigue damage of rotor components using National Instruments hardware and LabVIEW software. A full description of the data acquisition system including its configuration and salient features, is presented in this paper.

  4. Progress in Open Rotor Research: A U.S. Perspective

    NASA Technical Reports Server (NTRS)

    Van Zante, Dale E.

    2015-01-01

    In response to the 1970s oil crisis, NASA created the Advanced Turboprop Project (ATP) to mature technologies for high-speed propellers to enable large reductions in fuel burn relative to turbofan engines of that era. Both single rotation and contra- rotation concepts were designed and tested in ground based facilities as well as flight. Some novel concepts/configurations were proposed as part of the effort. The high-speed propeller concepts did provide fuel burn savings, albeit with some acoustics and structural challenges to overcome. When fuel prices fell, the business case for radical new engine configurations collapsed and the research emphasis returned to high bypass ducted configurations. With rising oil prices and increased environmental concerns there is renewed interest in high-speed propeller based engine architectures. Contemporary analysis tools for aerodynamics and aeroacoustics have enabled a new era of blade designs that have both high efficiency and lower noise characteristics. A recent series of tests in the U.S. have characterized the aerodynamic performance and noise from these modern contra-rotating propeller designs. Additionally the installation and noise shielding aspects for conventional airframes and blended wing bodies have been studied. Historical estimates of 'propfan' performance have relied on legacy propeller performance and acoustics data. Current system studies make use of the modern propeller data and higher fidelity installation effects data to estimate the performance of a contemporary aircraft system. Contemporary designs have demonstrated high net efficiency, approximately 86%, at 0.78 Mach, and low noise, greater than 15 EPNdB cumulative margin to Chapter 4 when analyzed on a NASA derived aircraft/mission. This paper presents the current state of high-speed propeller/open rotor research within the U.S. from an overall viewpoint of the various efforts ongoing. The remaining technical challenges to a production engine include

  5. Progress in Open Rotor Research: A U.S. Perspective

    NASA Technical Reports Server (NTRS)

    Van Zante, Dale E.

    2015-01-01

    In response to the 1970s oil crisis, NASA created the Advanced Turboprop Project (ATP) to mature technologies for high-speed propellers to enable large reductions in fuel burn relative to turbofan engines of that era. Both single rotation and contra-rotation concepts were designed and tested in ground based facilities as well as flight. Some novel concepts configurations that were not well publicized at the time, were proposed as part of the effort. The high-speed propeller concepts did provide fuel burn savings, albeit with some acoustics and structural challenges to overcome. When fuel prices fell, the business case for radical new engine configurations collapsed and the research emphasis returned to high bypass ducted configurations. With rising oil prices and increased environmental concerns there is renewed interest in high-speed propeller based engine architectures. Contemporary analysis tools for aerodynamics and aeroacoustics have enabled a new era of blade designs that have both high efficiency and acceptable noise characteristics. A recent series of tests in the U.S. have characterized the aerodynamic performance and noise from these modern contra-rotating propeller designs. Additionally the installation and noise shielding aspects for conventional airframes and blended wing bodies have been studied. Historical estimates of propfan performance have relied on legacy propeller performance and acoustics data. Current system studies make use of the modern propeller data with higher fidelity installation effects data to estimate the performance of a contemporary aircraft system with favorable results. This paper presents the current state of high-speed propeller open rotor research within the U.S. from an overall viewpoint of the various efforts ongoing. The current projections for the technology are presented.

  6. Numerical methods for problems in computational aeroacoustics

    NASA Astrophysics Data System (ADS)

    Mead, Jodi Lorraine

    1998-12-01

    A goal of computational aeroacoustics is the accurate calculation of noise from a jet in the far field. This work concerns the numerical aspects of accurately calculating acoustic waves over large distances and long time. More specifically, the stability, efficiency, accuracy, dispersion and dissipation in spatial discretizations, time stepping schemes, and absorbing boundaries for the direct solution of wave propagation problems are determined. Efficient finite difference methods developed by Tam and Webb, which minimize dispersion and dissipation, are commonly used for the spatial and temporal discretization. Alternatively, high order pseudospectral methods can be made more efficient by using the grid transformation introduced by Kosloff and Tal-Ezer. Work in this dissertation confirms that the grid transformation introduced by Kosloff and Tal-Ezer is not spectrally accurate because, in the limit, the grid transformation forces zero derivatives at the boundaries. If a small number of grid points are used, it is shown that approximations with the Chebyshev pseudospectral method with the Kosloff and Tal-Ezer grid transformation are as accurate as with the Chebyshev pseudospectral method. This result is based on the analysis of the phase and amplitude errors of these methods, and their use for the solution of a benchmark problem in computational aeroacoustics. For the grid transformed Chebyshev method with a small number of grid points it is, however, more appropriate to compare its accuracy with that of high- order finite difference methods. This comparison, for an order of accuracy 10-3 for a benchmark problem in computational aeroacoustics, is performed for the grid transformed Chebyshev method and the fourth order finite difference method of Tam. Solutions with the finite difference method are as accurate. and the finite difference method is more efficient than, the Chebyshev pseudospectral method with the grid transformation. The efficiency of the Chebyshev

  7. Design, fabrication, and testing of an ultrasonic de-icing system for helicopter rotor blades

    NASA Astrophysics Data System (ADS)

    Palacios, Jose Luis

    A low-power, non-thermal ultrasonic de-icing system is introduced as a possible substitute for current electro-thermal systems. The system generates delaminating ultrasonic transverse shear stresses at the interface of accreted ice. A PZT-4 disk driven at 28.5 KHz (radial resonance of the disk) instantaneously de-bonds 2 mm thick freezer ice layers. The ice layers are accreted to a 0.7 mm thick, 30.4 cm x 30.4 cm steel plate at an environment temperature of -20°C. A power input of 50 Watts is applied to the actuator (50 V, 19.6 KV/m), which translates to a de-icing power of 0.07 W/cm2. A finite element model of the actuator bonded to the isotropic plate is used to guide the design of the system, and predicts the transverse shear stresses at the ice interface. Wind tunnel icing tests were conducted to demonstrate the potential use of the proposed system under impact icing conditions. Both glaze ice and rime ice were generated on steel and composite plates by changing the cloud conditions of the wind tunnel. Continuous ultrasonic vibration prevented impact ice formation around the actuator location at an input power not exceeding 0.18 W/cm 2 (1.2 W/in2). As ice thickness reached a critical thickness of approximately 1.2 mm, shedding occurred on those locations where ultrasonic transverse shear stresses exceeded the shear adhesion strength of the ice. Finite element transverse shear stress predictions correlate with observed experimental impact ice de-bonding behavior. To increase the traveling distance of propagating ultrasonic waves, ultrasonic shear horizontal wave modes are studied. Wave modes providing large modal interface transverse shear stress concentration coefficients (ISCC) between the host structure (0.7 mm thick steel plate) and accreted ice (2.5 mm thick ice layer) are identified and investigated for a potential increase in the wave propagation distance. Ultrasonic actuators able to trigger these optimum wave modes are designed and fabricated. Despite

  8. Design and Use of Microphone Directional Arrays for Aeroacoustic Measurements

    NASA Technical Reports Server (NTRS)

    Humphreys, William M., Jr.; Brooks, Thomas F.; Hunter, William W., Jr.; Meadows, Kristine R.

    1998-01-01

    An overview of the development of two microphone directional arrays for aeroacoustic testing is presented. These arrays were specifically developed to measure airframe noise in the NASA Langley Quiet Flow Facility. A large aperture directional array using 35 flush-mounted microphones was constructed to obtain high resolution noise localization maps around airframe models. This array possesses a maximum diagonal aperture size of 34 inches. A unique logarithmic spiral layout design was chosen for the targeted frequency range of 2-30 kHz. Complementing the large array is a small aperture directional array, constructed to obtain spectra and directivity information from regions on the model. This array, possessing 33 microphones with a maximum diagonal aperture size of 7.76 inches, is easily moved about the model in elevation and azimuth. Custom microphone shading algorithms have been developed to provide a frequency- and position-invariant sensing area from 10-40 kHz with an overall targeted frequency range for the array of 5-60 kHz. Both arrays are employed in acoustic measurements of a 6 percent of full scale airframe model consisting of a main element NACA 632-215 wing section with a 30 percent chord half-span flap. Representative data obtained from these measurements is presented, along with details of the array calibration and data post-processing procedures.

  9. Aeroacoustic Analysis of Turbofan Noise Generation

    NASA Technical Reports Server (NTRS)

    Meyer, Harold D.; Envia, Edmane

    1996-01-01

    This report provides an updated version of analytical documentation for the V072 Rotor Wake/Stator Interaction Code. It presents the theoretical derivation of the equations used in the code and, where necessary, it documents the enhancements and changes made to the original code since its first release. V072 is a package of FORTRAN computer programs which calculate the in-duct acoustic modes excited by a fan/stator stage operating in a subsonic mean flow. Sound is generated by the stator vanes interacting with the mean wakes of the rotor blades. In this updated version, only the tonal noise produced at the blade passing frequency and its harmonics, is described. The broadband noise component analysis, which was part of the original report, is not included here. The code provides outputs of modal pressure and power amplitudes generated by the rotor-wake/stator interaction. The rotor/stator stage is modeled as an ensemble of blades and vanes of zero camber and thickness enclosed within an infinite hard-walled annular duct. The amplitude of each propagating mode is computed and summed to obtain the harmonics of sound power flux within the duct for both upstream and downstream propagating modes.

  10. Dynamical response of the Galileo Galilei on the ground rotor to test the equivalence principle: Theory, simulation, and experiment. II. The rejection of common mode forces

    NASA Astrophysics Data System (ADS)

    Comandi, G. L.; Toncelli, R.; Chiofalo, M. L.; Bramanti, D.; Nobili, A. M.

    2006-03-01

    "Galileo Galilei on the ground" (GGG) is a fast rotating differential accelerometer designed to test the equivalence principle (EP). Its sensitivity to differential effects, such as the effect of an EP violation, depends crucially on the capability of the accelerometer to reject all effects acting in common mode. By applying the theoretical and simulation methods reported in Part I of this work, and tested therein against experimental data, we predict the occurrence of an enhanced common mode rejection of the GGG accelerometer. We demonstrate that the best rejection of common mode disturbances can be tuned in a controlled way by varying the spin frequency of the GGG rotor.

  11. Dynamical response of the Galileo Galilei on the ground rotor to test the equivalence principle: Theory, simulation, and experiment. II. The rejection of common mode forces

    SciTech Connect

    Comandi, G.L.; Toncelli, R.; Chiofalo, M.L.; Bramanti, D.; Nobili, A.M.

    2006-03-15

    'Galileo Galilei on the ground' (GGG) is a fast rotating differential accelerometer designed to test the equivalence principle (EP). Its sensitivity to differential effects, such as the effect of an EP violation, depends crucially on the capability of the accelerometer to reject all effects acting in common mode. By applying the theoretical and simulation methods reported in Part I of this work, and tested therein against experimental data, we predict the occurrence of an enhanced common mode rejection of the GGG accelerometer. We demonstrate that the best rejection of common mode disturbances can be tuned in a controlled way by varying the spin frequency of the GGG rotor.

  12. Rotor/wing aerodynamic interactions in hover

    NASA Technical Reports Server (NTRS)

    Felker, F. F.; Light, J. S.

    1986-01-01

    An experimental and theoretical investigation of rotor/wing aerodynamic interactions in hover is described. The experimental investigation consisted of both a large-scale and small-scale test. A 0.658-scale, V-22 rotor and wing was used in the large-scale test. Wind download, wing surface pressure, rotor performance, and rotor downwash data from the large-scale test are presented. A small-scale experiment was conducted to determine how changes in the rotor/wing geometry affected the aerodynamic interactions. These geometry variations included the distance between the rotor and wing, wing incidence angle, and configurations both with the rotor axis at the tip of the wing (tilt rotor configuration) and with the rotor axis at the center of the wing (compound helicopter configuration). A wing with boundary-layer control was also tested to evaluate the effect of leading and trailing edge upper surface blowing on the wing download. A computationally efficient, semi-empirical theory was developed to predict the download on the wing. Finally, correlations between the theoretical predictions and test data are presented.

  13. Parametric tip effects for conformable rotor applications

    NASA Technical Reports Server (NTRS)

    Mantay, W. R.; Yeager, W. T., Jr.

    1983-01-01

    A research study was initiated to systematically determine the impact of selected blade tip geometric parameters on aeroelasticity conformable rotor performance and loads characteristics. The model articulated rotors included baseline and torsionally soft blades with interchangeable tips. Seven blade tip designs were evaluated on the baseline rotor and three tip designs were tested on the torsionally soft blades. The designs incorporated a systematic variation in three geometric parameters: sweep, taper, and anhedral. The rotors were evaluated in the NASA Langley Transonic Dynamics Tunnel at several advance ratios, lift and propulsive force values, and tip Mach numbers. Based on the test results, tip parameter variations generated significant rotor performance and loads difference for both baseline and torsionally soft blades. Azimuthal variation of elastic twist generated by the tip parameters strongly correlated with rotor performance and loads, but the magnitude of advancing blade elastic twist did not correlate.

  14. Computational Analysis of a Chevron Nozzle Uniquely Tailored for Propulsion Airframe Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Massey, Steven J.; Elmiligui, Alaa A.; Hunter, Craig A.; Thomas, Russell H.; Pao, S. Paul; Mengle, Vinod G.

    2006-01-01

    A computational flow field and predicted jet noise source analysis is presented for asymmetrical fan chevrons on a modern separate flow nozzle at take off conditions. The propulsion airframe aeroacoustic asymmetric fan nozzle is designed with an azimuthally varying chevron pattern with longer chevrons close to the pylon. A baseline round nozzle without chevrons and a reference nozzle with azimuthally uniform chevrons are also studied. The intent of the asymmetric fan chevron nozzle was to improve the noise reduction potential by creating a favorable propulsion airframe aeroacoustic interaction effect between the pylon and chevron nozzle. This favorable interaction and improved noise reduction was observed in model scale tests and flight test data and has been reported in other studies. The goal of this study was to identify the fundamental flow and noise source mechanisms. The flow simulation uses the asymptotically steady, compressible Reynolds averaged Navier-Stokes equations on a structured grid. Flow computations are performed using the parallel, multi-block, structured grid code PAB3D. Local noise sources were mapped and integrated computationally using the Jet3D code based upon the Lighthill Acoustic Analogy with anisotropic Reynolds stress modeling. In this study, trends of noise reduction were correctly predicted. Jet3D was also utilized to produce noise source maps that were then correlated to local flow features. The flow studies show that asymmetry of the longer fan chevrons near the pylon work to reduce the strength of the secondary flow induced by the pylon itself, such that the asymmetric merging of the fan and core shear layers is significantly delayed. The effect is to reduce the peak turbulence kinetic energy and shift it downstream, reducing overall noise production. This combined flow and noise prediction approach has yielded considerable understanding of the physics of a fan chevron nozzle designed to include propulsion airframe aeroacoustic

  15. Exploratory investigation of aeroacoustic optimization of the variable impedance edge concept applied to upper surface blown configurations

    NASA Technical Reports Server (NTRS)

    Hayden, R. E.

    1976-01-01

    The feasibility of using porous surfaces on lift augmentation flaps to reduce the noise at the source, without adversely affecting aerodynamic performance was investigated. Numerous flap configurations were tested on a USB (upper surface blowing) type powered lift model (approximately 1/5 full scale). Significant reductions of far field noise and aeroacoustic pressures were found, and many configurations exhibited aerodynamic characteristics comparable to unmodified configurations of the nozzle/flap system.

  16. Isolated Open Rotor Noise Prediction Assessment Using the F31A31 Historical Blade Set

    NASA Technical Reports Server (NTRS)

    Nark, Douglas M.; Jones, William T.; Boyd, D. Douglas, Jr.; Zawodny, Nikolas S.

    2016-01-01

    In an effort to mitigate next-generation fuel efficiency and environmental impact concerns for aviation, open rotor propulsion systems have received renewed interest. However, maintaining the high propulsive efficiency while simultaneously meeting noise goals has been one of the challenges in making open rotor propulsion a viable option. Improvements in prediction tools and design methodologies have opened the design space for next generation open rotor designs that satisfy these challenging objectives. As such, validation of aerodynamic and acoustic prediction tools has been an important aspect of open rotor research efforts. This paper describes validation efforts of a combined computational fluid dynamics and Ffowcs Williams and Hawkings equation methodology for open rotor aeroacoustic modeling. Performance and acoustic predictions were made for a benchmark open rotor blade set and compared with measurements over a range of rotor speeds and observer angles. Overall, the results indicate that the computational approach is acceptable for assessing low-noise open rotor designs. Additionally, this approach may be used to provide realistic incident source fields for acoustic shielding/scattering studies on various aircraft configurations.

  17. Forward velocity effects on fan noise and the influence of inlet aeroacoustic design as measured in the NASA Ames 40 x 80 foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Holm, R. G.; Langenbrunner, L. E.; Mccann, E. O.

    1981-01-01

    The inlet radiated noise of a turbofan engine was studied. The principal research objectives were to characterize or suppress such noise with particular regard to its tonal characteristics. The major portion of this research was conducted by using ground-based static testing without simulation of aircraft forward speed or aircraft installation-related aeroacoustic effects.

  18. Pneumatic boot for helicopter rotor deicing

    NASA Technical Reports Server (NTRS)

    Blaha, B. J.; Evanich, P. L.

    1981-01-01

    Pneumatic deicer boots for helicopter rotor blades were tested. The tests were conducted in the 6 by 9 ft icing research tunnel on a stationary section of a UH-IH helicopter main rotor blade. The boots were effective in removing ice and in reducing aerodynamic drag due to ice.

  19. Flywheel Rotor Safe-Life Technology

    NASA Technical Reports Server (NTRS)

    Ratner, J. K. H.; Chang, J. B.; Christopher, D. A.; McLallin, Kerry L. (Technical Monitor)

    2002-01-01

    Since the 1960s, research has been conducted into the use of flywheels as energy storage systems. The-proposed applications include energy storage for hybrid and electric automobiles, attitude control and energy storage for satellites, and uninterruptible power supplies for hospitals and computer centers. For many years, however, the use of flywheels for space applications was restricted by the total weight of a system employing a metal rotor. With recent technological advances in the manufacturing of composite materials, however, lightweight composite rotors have begun to be proposed for such applications. Flywheels with composite rotors provide much higher power and energy storage capabilities than conventional chemical batteries. However, the failure of a high speed flywheel rotor could be a catastrophic event. For this reason, flywheel rotors are classified by the NASA Fracture Control Requirements Standard as fracture critical parts. Currently, there is no industry standard to certify a composite rotor for safe and reliable operation forth( required lifetime of the flywheel. Technical problems hindering the development of this standard include composite manufacturing inconsistencies, insufficient nondestructive evaluation (NDE) techniques for detecting defects and/or impact damage, lack of standard material test methods for characterizing composite rotor design allowables, and no unified proof (over-spin) test for flight rotors. As part of a flywheel rotor safe-life certification pro-ram funded b the government, a review of the state of the art in composite rotors is in progress. The goal of the review is to provide a clear picture of composite flywheel rotor technologies. The literature review has concentrated on the following topics concerning composites and composite rotors: durability (fatigue) and damage tolerance (safe-life) analysis/test methods, in-service NDE and health monitoring techniques, spin test methods/ procedures, and containment options

  20. Analysis of rotor vibratory loads using higher harmonic pitch control

    NASA Technical Reports Server (NTRS)

    Quackenbush, Todd R.; Bliss, Donald B.; Boschitsch, Alexander H.; Wachspress, Daniel A.

    1992-01-01

    Experimental studies of isolated rotors in forward flight have indicated that higher harmonic pitch control can reduce rotor noise. These tests also show that such pitch inputs can generate substantial vibratory loads. The modification is summarized of the RotorCRAFT (Computation of Rotor Aerodynamics in Forward flighT) analysis of isolated rotors to study the vibratory loading generated by high frequency pitch inputs. The original RotorCRAFT code was developed for use in the computation of such loading, and uses a highly refined rotor wake model to facilitate this task. The extended version of RotorCRAFT incorporates a variety of new features including: arbitrary periodic root pitch control; computation of blade stresses and hub loads; improved modeling of near wake unsteady effects; and preliminary implementation of a coupled prediction of rotor airloads and noise. Correlation studies are carried out with existing blade stress and vibratory hub load data to assess the performance of the extended code.

  1. On aerodynamic design of the Savonius windmill rotor

    SciTech Connect

    Mojola, O.O.

    1982-08-01

    This paper examines under field conditions the performance characteristics of the Savonius windmill rotor. Test data were collected on the speed, torque and power of the rotor at a large number of wind speeds for each of seven values of the rotor overlap ratio. Field testing procedures are critically appraised and a unified approach is suggested. The performance data of the Savonius rotor are also fully discussed and design criteria established.

  2. Aeroacoustics of Turbulent High-Speed Jets

    NASA Technical Reports Server (NTRS)

    Rao, Ram Mohan; Lundgren, Thomas S.

    1996-01-01

    Aeroacoustic noise generation in a supersonic round jet is studied to understand in particular the effect of turbulence structure on the noise without numerically compromising the turbulence itself. This means that direct numerical simulations (DNS's) are needed. In order to use DNS at high enough Reynolds numbers to get sufficient turbulence structure we have decided to solve the temporal jet problem, using periodicity in the direction of the jet axis. Physically this means that turbulent structures in the jet are repeated in successive downstream cells instead of being gradually modified downstream into a jet plume. Therefore in order to answer some questions about the turbulence we will partially compromise the overall structure of the jet. The first section of chapter 1 describes some work on the linear stability of a supersonic round jet and the implications of this for the jet noise problem. In the second section we present preliminary work done using a TVD numerical scheme on a CM5. This work is only two-dimensional (plane) but shows very interesting results, including weak shock waves. However this is a nonviscous computation and the method resolves the shocks by adding extra numerical dissipation where the gradients are large. One wonders whether the extra dissipation would influence small turbulent structures like small intense vortices. The second chapter is an extensive discussion of preliminary numerical work using the spectral method to solve the compressible Navier-Stokes equations to study turbulent jet flows. The method uses Fourier expansions in the azimuthal and streamwise direction and a 1-D B-spline basis representation in the radial direction. The B-spline basis is locally supported and this ensures block diagonal matrix equations which are solved in O(N) steps. A very accurate highly resolved DNS of a turbulent jet flow is expected.

  3. Rotor Flapping Response to Active Control

    NASA Technical Reports Server (NTRS)

    Nguyen, Khanh; Johnson, Wayne

    2004-01-01

    Rotor active control using higher harmonic blade pitch has been proposed as a means to reduce both rotor radiated noise and airframe vibration and to enhance rotor performance. The higher harmonic input, however, can affect rotor thrust and cyclic flapping - the basic trim characteristics of the rotor. Some of the trim changes can negate the active control benefits. For example, wind tunnel test results of a full scale BO-105 rotor with individual-blade control indicate some rotor performance improvements, accompanied with changes in rotor trim, using two-per-rev blade pitch input. The observed performance benefits could therefore be a simple manifestation of the trim change rather than an efficient redistribution of the rotor airloads. More recently, the flight test of the BO-105 helicopter equip,ped with individual-blade-control actuators also reported trim changes whenever the two-per-rev blade pitch for noise reduction was activated. The pilot had to adjust the trim control to maintain the aircraft under a constant flight path. These two cases highlight the, importance of trim considerations in the application of active control to rotorcraft.

  4. Optimal Aeroacoustic Shape Design Using the Surrogate Management Framework

    DTIC Science & Technology

    2004-02-09

    wish to thank the IMA for providing a forum for collaboration, as well as Charles Audet and Petros Koumoutsakos for valuable discussions. The authors...17] N. Hansen, D. Mller, and P. Koumoutsakos . Reducing the time complexity of the derandomized evolution strategy with covariance matrix adaptation...P. Koumoutsakos . Optimal aeroacoustic shape design using approximation modeling. Annual Research Briefs, Center for Turbulence Research, Stanford

  5. Second Computational Aeroacoustics (CAA) Workshop on Benchmark Problems

    NASA Technical Reports Server (NTRS)

    Tam, C. K. W. (Editor); Hardin, J. C. (Editor)

    1997-01-01

    The proceedings of the Second Computational Aeroacoustics (CAA) Workshop on Benchmark Problems held at Florida State University are the subject of this report. For this workshop, problems arising in typical industrial applications of CAA were chosen. Comparisons between numerical solutions and exact solutions are presented where possible.

  6. On least-order flow decompositions for aerodynamics and aeroacoustics

    NASA Astrophysics Data System (ADS)

    Schlegel, Michael; Noack, Bernd R.; Jordan, Peter

    2012-11-01

    A generalisation of proper orthogonal decomposition (POD) for optimal flow resolution of linearly related observables is presented, as proposed in the identically named publication of Schlegel, Noack, Jordan, Dillmann, Groeschel, Schroeder, Wei, Freund, Lehmann and Tadmor (Journal of Fluid Mechanics 2012, vol. 697, pp. 367-398). This Galerkin expansion, termed ``observable inferred decomposition'' (OID), addresses a need in aerodynamic and aeroacoustic applications by identifying the modes contributing most to these observables. Thus, OID constitutes a building block for physical understanding, least-biased conditional sampling, state estimation and control design. From a continuum of OID versions, two variants are tailored for purposes of observer and control design, respectively. Three aerodynamic and aeroacoustic observables are studied: (1) lift and drag fluctuation of a two-dimensional cylinder wake flow, (2) aeroacoustic density fluctuations measured by a sensor array and emitted from a two-dimensional compressible mixing layer, and (3) aeroacoustic pressure monitored by a sensor array and emitted from a three-dimensional compressible jet. The most ``drag-related,'' ``lift-related'' and ``loud'' structures are distilled and interpreted in terms of known physical processes. This work was partially funded by the DFG under grants SCHL 586/2-1 and ANR, Chair of Excellence, TUCOROM.

  7. Flapping inertia for selected rotor blades

    NASA Technical Reports Server (NTRS)

    Berry, John D.; May, Matthew J.

    1991-01-01

    Aerodynamics of helicopter rotor systems cannot be investigated without consideration for the dynamics of the rotor. One of the principal properties of the rotor which affects the rotor dynamics is the inertia of the rotor blade about its root attachment. Previous aerodynamic investigation have been performed on rotor blades with a variety of planforms to determine the performance differences due to blade planform. The blades tested for this investigation have been tested on the U.S. Army 2 meter rotor test system (2MRTS) in the NASA Langley 14 by 22 foot subsonic tunnel for hover performance. This investigation was intended to provide fundamental information on the flapping inertia of five rotor blades with differing planforms. The inertia of the bare cuff and the cuff with a blade extension were also measured for comparison with the inertia of the blades. Inertia was determined using a swing testing technique, using the period of oscillation to determine the effective flapping inertia. The effect of damping in the swing test was measured and described. A comparison of the flapping inertials for rectangular and tapered planform blades of approximately the same mass showed the tapered blades to have a lower inertia, as expected.

  8. Aeroacoustic characterization of scaled canonical nose landing gear configurations

    NASA Astrophysics Data System (ADS)

    Zawodny, Nikolas S.

    Aircraft noise is a critical issue in the commercial airline industry. Airframe noise is a subcomponent of aircraft noise and is generally dominant over jet engine noise during approach conditions, which can lead to high community impact. Landing gears have been identified as major components of airframe noise during landing configurations for commercial aircraft. They are perhaps the least understood contributors to airframe noise due to complex flow patterns associated with intricate gear component geometries. Nose landing gear in particular have received much attention in recent years, exhibiting acoustic signatures on the order of the main landing gear assembly of an aircraft, while simultaneously being more amenable to scaled wind tunnel testing. In order to characterize the acoustic signature of a complex geometry such as a nose landing gear, it is important to isolate, study, and understand the acoustic contributions of individual component geometries. The purpose of this dissertation is to develop a correlation between the complex flow field nature and far-field acoustic signature of a nose landing gear sub-system. The model under investigation is a 1/2-scale shock-strut cylinder coupled with an adjustable torque link apparatus. This geometry was chosen due to its fundamental importance and implementation across a wide span of commercial aircraft. The fluid dynamic (surface pressure and stereoscopic particle image velocimety) and aeroacoustic (far-field microphone and phased array) experiments were performed in the University of Florida Aeroacoustic Flow Facility. The experimental data compare favorably with the results of a numerical simulation using PowerFLOW, a lattice-Boltzmann solver developed by the Exa Corporation. The far-field acoustic results of this dissertation have shown non-uniform scaling behavior as a function of frequency for the different model configurations tested. For frequencies that appropriately satisfied the condition of acoustic

  9. Development of Improved Surface Integral Methods for Jet Aeroacoustic Predictions

    NASA Technical Reports Server (NTRS)

    Pilon, Anthony R.; Lyrintzis, Anastasios S.

    1997-01-01

    The accurate prediction of aerodynamically generated noise has become an important goal over the past decade. Aeroacoustics must now be an integral part of the aircraft design process. The direct calculation of aerodynamically generated noise with CFD-like algorithms is plausible. However, large computer time and memory requirements often make these predictions impractical. It is therefore necessary to separate the aeroacoustics problem into two parts, one in which aerodynamic sound sources are determined, and another in which the propagating sound is calculated. This idea is applied in acoustic analogy methods. However, in the acoustic analogy, the determination of far-field sound requires the solution of a volume integral. This volume integration again leads to impractical computer requirements. An alternative to the volume integrations can be found in the Kirchhoff method. In this method, Green's theorem for the linear wave equation is used to determine sound propagation based on quantities on a surface surrounding the source region. The change from volume to surface integrals represents a tremendous savings in the computer resources required for an accurate prediction. This work is concerned with the development of enhancements of the Kirchhoff method for use in a wide variety of aeroacoustics problems. This enhanced method, the modified Kirchhoff method, is shown to be a Green's function solution of Lighthill's equation. It is also shown rigorously to be identical to the methods of Ffowcs Williams and Hawkings. This allows for development of versatile computer codes which can easily alternate between the different Kirchhoff and Ffowcs Williams-Hawkings formulations, using the most appropriate method for the problem at hand. The modified Kirchhoff method is developed primarily for use in jet aeroacoustics predictions. Applications of the method are shown for two dimensional and three dimensional jet flows. Additionally, the enhancements are generalized so that

  10. A study of the noise radiation from four helicopter rotor blades. [tests in Ames 40 by 20 foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Lee, A.; Mosher, M.

    1978-01-01

    Acoustic measurements were taken of a modern helicopter rotor with four blade tip shapes in the NASA Ames 40-by-80-Foot Wind Tunnel. The four tip shapes are: rectangular, swept, trapezoidal, and swept tapered in platform. Acoustic effects due to tip shape changes were studied based on the dBA level, peak noise pressure, and subjective rating. The swept tapered blade was found to be the quietest above an advancing tip Mach number of about 0.9, and the swept blade was the quietest at low speed. The measured high speed impulsive noise was compared with theoretical predictions based on thickness effects; good agreement was found.

  11. Theoretical models of helicopter rotor noise

    NASA Technical Reports Server (NTRS)

    Hawkings, D. L.

    1978-01-01

    For low speed rotors, it is shown that unsteady load models are only partially successful in predicting experimental levels. A theoretical model is presented which leads to the concept of unsteady thickness noise. This gives better agreement with test results. For high speed rotors, it is argued that present models are incomplete and that other mechanisms are at work. Some possibilities are briefly discussed.

  12. Rotor systems research aircraft simulation mathematical model

    NASA Technical Reports Server (NTRS)

    Houck, J. A.; Moore, F. L.; Howlett, J. J.; Pollock, K. S.; Browne, M. M.

    1977-01-01

    An analytical model developed for evaluating and verifying advanced rotor concepts is discussed. The model was used during in both open loop and real time man-in-the-loop simulation during the rotor systems research aircraft design. Future applications include: pilot training, preflight of test programs, and the evaluation of promising concepts before their implementation on the flight vehicle.

  13. Laser balancing demonstration on a high-speed flexible rotor

    NASA Technical Reports Server (NTRS)

    Demuth, R. S.; Rio, R. A.; Fleming, D. P.

    1979-01-01

    This paper describes a flexible rotor system used for two-plane laser balancing and an experimental demonstration of the laser material removal method for balancing. A laboratory test rotor was modified to accept balancing corrections using a laser metal removal method while the rotor is at operating speed. The laser setup hardware required to balance the rotor using two correction planes is described. The test rig optical configuration and a neodymium glass laser were assembled and calibrated for material removal rates. Rotor amplitudes before and after balancing, trial and correction weights, rotor speed during operation of laser, and balancing time were documented. The rotor was balanced through the first bending critical speed using the laser material removal procedure to apply trial weights and correction weights without stopping the rotor.

  14. Tip-path-plane angle effects on rotor blade-vortex interaction noise levels and directivity

    NASA Technical Reports Server (NTRS)

    Burley, Casey L.; Martin, Ruth M.

    1988-01-01

    Acoustic data of a scale model BO-105 main rotor acquired in a large aeroacoustic wind tunnel are presented to investigate the parametric effects of rotor operating conditions on blade-vortex interaction (BVI) impulsive noise. Contours of a BVI noise metric are employed to quantify the effects of rotor advance ratio and tip-path-plane angle on BVI noise directivity and amplitude. Acoustic time history data are presented to illustrate the variations in impulsive characteristics. The directionality, noise levels and impulsive content of both advancing and retreating side BVI are shown to vary significantly with tip-path-plane angle and advance ratio over the range of low and moderate flight speeds considered.

  15. Measurement of aeroacoustic noise generated on wind turbine blades modified by trailing edge brushes

    NASA Astrophysics Data System (ADS)

    Asheim, Michael J.

    As wind technology becomes a larger portion of the energy production picture, the problematic interactions between the machines and society will continue to become more pronounced. Of these problems, wind turbine noise is one of the most important to the future of wind turbine development. This study looks at the effect trailing edge brushes mounted on the 2 bladed Controls Advance Research Turbine (CART 2), located at the National Wind Technology Center, have on the overall acoustic and aerodynamic performance of the blades. The use of trailing edge brushes reduced the aeroacoustic noise by 1.0 to 5.0 dB over the baseline blade, depending on wind speed. This acoustic performance comes at a cost to the aerodynamic performance of the blades. The aerodynamic performance indicators, such as turbine power and root bending moments show that increased drag due to the brushes is the main contributor to the reduction in power production. An economic analysis also investigated how to best use noise mitigation devices to optimize acoustic, power performance and loads of a 600 kW baseline turbine, such as the CART 2. The analysis shows that the use of up a noise mitigation device of 4 dB is best used by increasing the rotor diameter and the power rating of the machine, from a 43.3 m diameter, 600 kW machine to a 68.8 m diameter, 886.7 kW machine. This increase resulted in an annual energy production increase of 414% when using a Rayleigh wind distribution with at a mean annual wind speed of 8.5 m/s. This is a reduction of cost of energy from 0.0463 per kWh to 0.0422 kWh. This reduction in energy production costs helps to explain the continuing trend of turbine machine growth in both rotor diameter and power rating.

  16. A multi-fidelity framework for physics based rotor blade simulation and optimization

    NASA Astrophysics Data System (ADS)

    Collins, Kyle Brian

    with lower fidelity models. This thesis documents the development of automated low and high fidelity physics based rotor simulation frameworks. The low fidelity framework uses a comprehensive code with simplified aerodynamics. The high fidelity model uses a parallel processor capable CFD/CSD methodology. Both low and high fidelity frameworks include an aeroacoustic simulation for prediction of noise. A synergistic process is developed that uses both the low and high fidelity frameworks together to build approximate models of important high fidelity metrics as functions of certain design variables. To test the process, a 4-bladed hingeless rotor model is used as a baseline. The design variables investigated include tip geometry and spanwise twist distribution. Approximation models are built for metrics related to rotor efficiency and vibration using the results from 60+ high fidelity (CFD/CSD) experiments and 400+ low fidelity experiments. Optimization using the approximation models found the Pareto Frontier anchor points, or the design having maximum rotor efficiency and the design having minimum vibration. Various Pareto generation methods are used to find designs on the frontier between these two anchor designs. When tested in the high fidelity framework, the Pareto anchor designs are shown to be very good designs when compared with other designs from the high fidelity database. This provides evidence that the process proposed has merit. Ultimately, this process can be utilized by industry rotor designers with their existing tools to bring high fidelity analysis into the preliminary design stage of rotors. In conclusion, the methods developed and documented in this thesis have made several novel contributions. First, an automated high fidelity CFD based forward flight simulation framework has been built for use in preliminary design optimization. The framework was built around an integrated, parallel processor capable CFD/CSD/AA process. Second, a novel method of

  17. Internal rotor friction instability

    NASA Technical Reports Server (NTRS)

    Walton, J.; Artiles, A.; Lund, J.; Dill, J.; Zorzi, E.

    1990-01-01

    The analytical developments and experimental investigations performed in assessing the effect of internal friction on rotor systems dynamic performance are documented. Analytical component models for axial splines, Curvic splines, and interference fit joints commonly found in modern high speed turbomachinery were developed. Rotor systems operating above a bending critical speed were shown to exhibit unstable subsynchronous vibrations at the first natural frequency. The effect of speed, bearing stiffness, joint stiffness, external damping, torque, and coefficient of friction, was evaluated. Testing included material coefficient of friction evaluations, component joint quantity and form of damping determinations, and rotordynamic stability assessments. Under conditions similar to those in the SSME turbopumps, material interfaces experienced a coefficient of friction of approx. 0.2 for lubricated and 0.8 for unlubricated conditions. The damping observed in the component joints displayed nearly linear behavior with increasing amplitude. Thus, the measured damping, as a function of amplitude, is not represented by either linear or Coulomb friction damper models. Rotordynamic testing of an axial spline joint under 5000 in.-lb of static torque, demonstrated the presence of an extremely severe instability when the rotor was operated above its first flexible natural frequency. The presence of this instability was predicted by nonlinear rotordynamic time-transient analysis using the nonlinear component model developed under this program. Corresponding rotordynamic testing of a shaft with an interference fit joint demonstrated the presence of subsynchronous vibrations at the first natural frequency. While subsynchronous vibrations were observed, they were bounded and significantly lower in amplitude than the synchronous vibrations.

  18. Predesign study for a modern 4-bladed rotor for the NASA rotor systems research aircraft

    NASA Technical Reports Server (NTRS)

    Bishop, H. E.; Burkam, J. E.; Heminway, R. C.; Keys, C. N.; Smith, K. E.; Smith, J. H.; Staley, J. A.

    1981-01-01

    Trade-off study results and the rationale for the final selection of an existing modern four-bladed rotor system that can be adapted for installation on the Rotor Systems Research Aircraft (RSRA) are reported. The results of the detailed integration studies, parameter change studies, and instrumentation studies and the recommended plan for development and qualification of the rotor system is also given. Its parameter variants, integration on the RSRA, and support of ground and flight test programs are also discussed.

  19. Supersonic Quadrupole Noise Theory for High-Speed Helicopter Rotors

    NASA Technical Reports Server (NTRS)

    Farassat, F.; Brentner, Kenneth S.

    1997-01-01

    High-speed helicopter rotor impulsive noise prediction is an important problem of aeroacoustics. The deterministic quadrupoles have been shown to contribute significantly to high-speed impulsive (HSI) noise of rotors, particularly when the phenomenon of delocalization occurs. At high rotor-tip speeds, some of the quadrupole sources lie outside the sonic circle and move at supersonic speed. Brentner has given a formulation suitable for efficient prediction of quadrupole noise inside the sonic circle. In this paper, we give a simple formulation based on the acoustic analogy that is valid for both subsonic and supersonic quadrupole noise prediction. Like the formulation of Brentner, the model is exact for an observer in the far field and in the rotor plane and is approximate elsewhere. We give the full analytic derivation of this formulation in the paper. We present the method of implementation on a computer for supersonic quadrupoles using marching cubes for constructing the influence surface (Sigma surface) of an observer space- time variable (x; t). We then present several examples of noise prediction for both subsonic and supersonic quadrupoles. It is shown that in the case of transonic flow over rotor blades, the inclusion of the supersonic quadrupoles improves the prediction of the acoustic pressure signature. We show the equivalence of the new formulation to that of Brentner for subsonic quadrupoles. It is shown that the regions of high quadrupole source strength are primarily produced by the shock surface and the flow over the leading edge of the rotor. The primary role of the supersonic quadrupoles is to increase the width of a strong acoustic signal.

  20. Open Rotor - Analysis of Diagnostic Data

    NASA Technical Reports Server (NTRS)

    Envia, Edmane

    2011-01-01

    NASA is researching open rotor propulsion as part of its technology research and development plan for addressing the subsonic transport aircraft noise, emission and fuel burn goals. The low-speed wind tunnel test for investigating the aerodynamic and acoustic performance of a benchmark blade set at the approach and takeoff conditions has recently concluded. A high-speed wind tunnel diagnostic test campaign has begun to investigate the performance of this benchmark open rotor blade set at the cruise condition. Databases from both speed regimes will comprise a comprehensive collection of benchmark open rotor data for use in assessing/validating aerodynamic and noise prediction tools (component & system level) as well as providing insights into the physics of open rotors to help guide the development of quieter open rotors.

  1. Acoustic testing of a supersonic tip speed fan with acoustic treatment and rotor casting slots. Quiet engine program scale model fan C

    NASA Technical Reports Server (NTRS)

    Kazin, S. B.

    1973-01-01

    Acoustic tests were conducted on a high tip speed (1550 ft/sec, 472.44 m/sec) single stage fan with varying amounts of wall acoustic treatment and with circumferential slots over the rotor blade tips. The slots were also tested with acoustic treatment placed behind the slots. The wall treatment results show that the inlet treatment is more effective at high fan speeds and aft duct treatment is more effective at low fan speeds. Maximum PNL's on a 200-foot (60.96 m) sideline show the untreated slots to have increased the rear radiated noise at approach. However, when the treatment was added to the slots inlet radiated noise was decreased, resulting in little change relative to the solid casing on an EPNL basis.

  2. Experimental study of main rotor tip geometry and tail rotor interactions in hover. Volume 2: Run log and tabulated data

    NASA Technical Reports Server (NTRS)

    Balch, D. T.; Lombardi, J.

    1985-01-01

    A model scale hover test was conducted in the Sikorsky Aircraft Model Rotor hover Facility to identify and quantify the impact of the tail rotor on the demonstrated advantages of advanced geometry tip configurations. The existence of mutual interference between hovering main rotor and a tail rotor was acknowledged in the test. The test was conducted using the Basic Model Test Rig and two scaled main rotor systems, one representing a 1/5.727 scale UH-60A BLACK HAWK and the others a 1/4.71 scale S-76. Eight alternate rotor tip configurations were tested, 3 on the BLACK HAWK rotor and 6 on the S-76 rotor. Four of these tips were then selected for testing in close proximity to an operating tail rotor (operating in both tractor and pusher modes) to determine if the performance advantages that could be obtained from the use of advanced geometry tips in a main rotor only environment would still exist in the more complex flow field involving a tail rotor. This volume contains the test run log and tabulated data.

  3. Tip Vortex and Wake Characteristics of a Counterrotating Open Rotor

    NASA Technical Reports Server (NTRS)

    VanZante, Dale E.; Wernet, Mark P.

    2012-01-01

    One of the primary noise sources for Open Rotor systems is the interaction of the forward rotor tip vortex and blade wake with the aft rotor. NASA has collaborated with General Electric on the testing of a new generation of low noise, counterrotating Open Rotor systems. Three-dimensional particle image velocimetry measurements were acquired in the intra-rotor gap of the Historical Baseline blade set. The velocity measurements are of sufficient resolution to characterize the tip vortex size and trajectory as well as the rotor wake decay and turbulence character. The tip clearance vortex trajectory is compared to results from previously developed models. Forward rotor wake velocity profiles are shown. Results are presented in a form as to assist numerical modeling of Open Rotor system aerodynamics and acoustics.

  4. Measurements of atmospheric turbulence effects on tail rotor acoustics

    NASA Technical Reports Server (NTRS)

    Hagen, Martin J.; Yamauchi, Gloria K.; Signor, David B.; Mosher, Marianne

    1994-01-01

    Results from an outdoor hover test of a full-scale Lynx tail rotor are presented. The investigation was designed to further the understanding of the acoustics of an isolated tail rotor hovering out-of-ground effect in atmospheric turbulence, without the effects of the main rotor wake or other helicopter components. Measurements include simultaneous rotor performance, noise, inflow, and far-field atmospheric turbulence. Results with grid-generated inflow turbulence are also presented. The effects of atmospheric turbulence ingestion on rotor noise are quantified. In contradiction to current theories, increasing rotor inflow and rotor thrust were found to increase turbulence ingestion noise. This is the final report of Task 13A--Helicopter Tail Rotor Noise, of the NASA/United Kingdom Defense Research Agency cooperative Aeronautics Research Program.

  5. Optimizing tuning masses for helicopter rotor blade vibration reduction including computed airloads and comparison with test data

    NASA Technical Reports Server (NTRS)

    Pritchard, Jocelyn I.; Adelman, Howard M.; Walsh, Joanne L.; Wilbur, Matthew L.

    1992-01-01

    An optimization procedure is developed for locating tuning masses on a rotor blade so that vibratory loads are minimized and hub-shear harmonics are reduced without adding a large mass penalty. The airloads are computed by means of a helicopter analysis for the cases of three vs six tuning masses, with attention given to the prediction of changes in airloads. Frequencies, airloads, and hub loads are computed with the CAMRAD/JA helicopter analysis code and the Conmin general-purpose optimization program. The hub shear is found to be significantly reduced in both cases with the added mass, and the reduction of hub shear is demonstrated under three flight conditions. Comparisons with wind-tunnel data demonstrate that the correlation of mass location is good and the relationship between mass location and flight speed is predicted well by the model.

  6. Internal Friction And Instabilities Of Rotors

    NASA Technical Reports Server (NTRS)

    Walton, J.; Artiles, A.; Lund, J.; Dill, J.; Zorzi, E.

    1992-01-01

    Report describes study of effects of internal friction on dynamics of rotors prompted by concern over instabilities in rotors of turbomachines. Theoretical and experimental studies described. Theoretical involved development of nonlinear mathematical models of internal friction in three joints found in turbomachinery - axial splines, Curvic(TM) splines, and interference fits between smooth cylindrical surfaces. Experimental included traction tests to determine the coefficients of friction of rotor alloys at various temperatures, bending-mode-vibration tests of shafts equipped with various joints and rotordynamic tests of shafts with axial-spline and interference-fit joints.

  7. On Cup Anemometer Rotor Aerodynamics

    PubMed Central

    Pindado, Santiago; Pérez, Javier; Avila-Sanchez, Sergio

    2012-01-01

    The influence of anemometer rotor shape parameters, such as the cups' front area or their center rotation radius on the anemometer's performance was analyzed. This analysis was based on calibrations performed on two different anemometers (one based on magnet system output signal, and the other one based on an opto-electronic system output signal), tested with 21 different rotors. The results were compared to the ones resulting from classical analytical models. The results clearly showed a linear dependency of both calibration constants, the slope and the offset, on the cups' center rotation radius, the influence of the front area of the cups also being observed. The analytical model of Kondo et al. was proved to be accurate if it is based on precise data related to the aerodynamic behavior of a rotor's cup. PMID:22778638

  8. On cup anemometer rotor aerodynamics.

    PubMed

    Pindado, Santiago; Pérez, Javier; Avila-Sanchez, Sergio

    2012-01-01

    The influence of anemometer rotor shape parameters, such as the cups' front area or their center rotation radius on the anemometer's performance was analyzed. This analysis was based on calibrations performed on two different anemometers (one based on magnet system output signal, and the other one based on an opto-electronic system output signal), tested with 21 different rotors. The results were compared to the ones resulting from classical analytical models. The results clearly showed a linear dependency of both calibration constants, the slope and the offset, on the cups' center rotation radius, the influence of the front area of the cups also being observed. The analytical model of Kondo et al. was proved to be accurate if it is based on precise data related to the aerodynamic behavior of a rotor's cup.

  9. Structural integrity design for an active helicopter rotor blade with piezoelectric flap actuators

    NASA Astrophysics Data System (ADS)

    Lee, Jaehwan; Shin, SangJoon

    2011-04-01

    Helicopter uses a rotor system to generate lift, thrust and forces, and its aerodynamic environment is generally complex. Unsteady aerodynamic environment arises such as blade vortex interaction. This unsteady aerodynamic environment induces vibratory aerodynamic loads and high aeroacoustic noise. The aerodynamic load and aeroacoustic noise is at N times the rotor blade revolutions (N/rev). But conventional rotor control system composed of pitch links and swash plate is not capable of adjusting such vibratory loads because its control is restricted to 1/rev. Many active control methodologies have been examined to alleviate the problem. The blade using active control device manipulates the blade pitch angle with N/rev. In this paper, Active Trailing-edge Flap blade, which is one of the active control methods, is designed to reduce the unsteady aerodynamic loads. Active Trailing-edge Flap blade uses a trailing edge flap manipulated by an actuator to change camber line of the airfoil. Piezoelectric actuators are installed inside the blade to manipulate the trailing edge flap.

  10. Preliminary simulation of an advanced, hingless rotor XV-15 tilt-rotor aircraft

    NASA Technical Reports Server (NTRS)

    Mcveigh, M. A.

    1976-01-01

    The feasibility of the tilt-rotor concept was verified through investigation of the performance, stability and handling qualities of the XV-15 tilt rotor. The rotors were replaced by advanced-technology fiberglass/composite hingless rotors of larger diameter, combined with an advanced integrated fly-by-wire control system. A parametric simulation model of the HRXV-15 was developed, model was used to define acceptable preliminary ranges of primary and secondary control schedules as functions of the flight parameters, to evaluate performance, flying qualities and structural loads, and to have a Boeing-Vertol pilot conduct a simulated flight test evaluation of the aircraft.

  11. Domain decomposition for aerodynamic and aeroacoustic analyses, and optimization

    NASA Technical Reports Server (NTRS)

    Baysal, Oktay

    1995-01-01

    The overarching theme was the domain decomposition, which intended to improve the numerical solution technique for the partial differential equations at hand; in the present study, those that governed either the fluid flow, or the aeroacoustic wave propagation, or the sensitivity analysis for a gradient-based optimization. The role of the domain decomposition extended beyond the original impetus of discretizing geometrical complex regions or writing modular software for distributed-hardware computers. It induced function-space decompositions and operator decompositions that offered the valuable property of near independence of operator evaluation tasks. The objectives have gravitated about the extensions and implementations of either the previously developed or concurrently being developed methodologies: (1) aerodynamic sensitivity analysis with domain decomposition (SADD); (2) computational aeroacoustics of cavities; and (3) dynamic, multibody computational fluid dynamics using unstructured meshes.

  12. Advances in Numerical Boundary Conditions for Computational Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Tam, Christopher K. W.

    1997-01-01

    Advances in Computational Aeroacoustics (CAA) depend critically on the availability of accurate, nondispersive, least dissipative computation algorithm as well as high quality numerical boundary treatments. This paper focuses on the recent developments of numerical boundary conditions. In a typical CAA problem, one often encounters two types of boundaries. Because a finite computation domain is used, there are external boundaries. On the external boundaries, boundary conditions simulating the solution outside the computation domain are to be imposed. Inside the computation domain, there may be internal boundaries. On these internal boundaries, boundary conditions simulating the presence of an object or surface with specific acoustic characteristics are to be applied. Numerical boundary conditions, both external or internal, developed for simple model problems are reviewed and examined. Numerical boundary conditions for real aeroacoustic problems are also discussed through specific examples. The paper concludes with a description of some much needed research in numerical boundary conditions for CAA.

  13. Constrained Aeroacoustic Shape Optimization Using the Surrogate Management Framework

    NASA Technical Reports Server (NTRS)

    Marsden, Alison L.; Wang, Meng; Dennis, John E., Jr.

    2003-01-01

    Reduction of noise generated by turbulent flow past the trailing-edge of a lifting surface is a challenge in many aeronautical and naval applications. Numerical predictions of trailing-edge noise necessitate the use of advanced simulation techniques such as large-eddy simulation (LES) in order to capture a wide range of turbulence scales which are the source of broadband noise. Aeroacoustic calculations of the flow over a model airfoil trailing edge using LES and aeroacoustic theory have been presented in Wang and Moin and were shown to agree favorably with experiments. The goal of the present work is to apply shape optimization to the trailing edge flow previously studied, in order to control aerodynamic noise.

  14. Third Computational Aeroacoustics (CAA) Workshop on Benchmark Problems

    NASA Technical Reports Server (NTRS)

    Dahl, Milo D. (Editor)

    2000-01-01

    The proceedings of the Third Computational Aeroacoustics (CAA) Workshop on Benchmark Problems cosponsored by the Ohio Aerospace Institute and the NASA Glenn Research Center are the subject of this report. Fan noise was the chosen theme for this workshop with representative problems encompassing four of the six benchmark problem categories. The other two categories were related to jet noise and cavity noise. For the first time in this series of workshops, the computational results for the cavity noise problem were compared to experimental data. All the other problems had exact solutions, which are included in this report. The Workshop included a panel discussion by representatives of industry. The participants gave their views on the status of applying computational aeroacoustics to solve practical industry related problems and what issues need to be addressed to make CAA a robust design tool.

  15. Aeroacoustics of a porous plug jet noise suppressor

    NASA Technical Reports Server (NTRS)

    Dosanjh, D. S.

    1981-01-01

    The aeroacoustics of a porous plug jet noise suppressor was investigated. The predicted flow features of isentropic plug nozzles for different pressure ratios or exit flow Mach numbers, throat areas, ratios of the plug to annular nozzle radii, mass flow rates and the available run times possible with the existing compressed air supply system, are compiled. The dimensions and the coordinates of the contour of typical isentropic external expansion plugs with different exit flow Mach numbers are listed. Design details of the experimental facility and the plug nozzle selected for experimental aeroacoustic studies are reported. The analytical flow prediction by method of characteristics of a conical porous plug nozzles is initiated. The role of the shape, size, and porosity of the plug surface in achieving over a perforated conical plug a nearly isentropic shockfree supersonic flow field which is closely similar to the flow field of a contoured isentropic plug nozzle is examined.

  16. Numerical analysis of a variable camber rotor blade as a lift control device

    NASA Technical Reports Server (NTRS)

    Awani, A. O.; Stroub, R. H.

    1984-01-01

    A new rotor configuration called the variable camber rotor was numerically investigated as a lift control device. This rotor differs from a conventional (baseline) rotor only in the blade aft section. In this configuration, the aft section or flap is attached to the forward section by pin joint arrangement, and also connected to the rotor control system for the control of rotor thrust level and vectoring. Pilot action to the flap deflection controls rotor lift and tip path plane tilt. The drag due to flaps is presented and the theoretical result correlated with test data. The assessment of payoff for the variable camber rotor in comparison with conventional (baseline) rotor was examined in hover. The variable camber rotor is shown to increase hover power required by 1.35%, but such a minimal power penalty is not significant enough to be considered a negative result. In forward flight, the control needs of the variable camber rotor were evaluated.

  17. Adaptive beamforming for array signal processing in aeroacoustic measurements.

    PubMed

    Huang, Xun; Bai, Long; Vinogradov, Igor; Peers, Edward

    2012-03-01

    Phased microphone arrays have become an important tool in the localization of noise sources for aeroacoustic applications. In most practical aerospace cases the conventional beamforming algorithm of the delay-and-sum type has been adopted. Conventional beamforming cannot take advantage of knowledge of the noise field, and thus has poorer resolution in the presence of noise and interference. Adaptive beamforming has been used for more than three decades to address these issues and has already achieved various degrees of success in areas of communication and sonar. In this work an adaptive beamforming algorithm designed specifically for aeroacoustic applications is discussed and applied to practical experimental data. It shows that the adaptive beamforming method could save significant amounts of post-processing time for a deconvolution method. For example, the adaptive beamforming method is able to reduce the DAMAS computation time by at least 60% for the practical case considered in this work. Therefore, adaptive beamforming can be considered as a promising signal processing method for aeroacoustic measurements.

  18. Aeroacoustics of Propulsion Airframe Integration: Overview of NASA's Research

    NASA Technical Reports Server (NTRS)

    Thomas, Russell H.

    2003-01-01

    The integration of propulsion and airframe is fundamental to the design of an aircraft system. Many considerations influence the integration, such as structural, aerodynamic, and maintenance factors. In regard to the acoustics of an aircraft, the integration can have significant effects on the net radiated noise. Whether an engine is mounted above a wing or below can have a significant effect on noise that reaches communities below because of shielding or reflection of engine noise. This is an obvious example of the acoustic effects of propulsion airframe installation. Another example could be the effect of the pylon on the development of the exhaust plume and on the resulting jet noise. In addition, for effective system noise reduction the impact that installation has on noise reduction devices developed on isolated components must be understood. In the future, a focus on the aerodynamic and acoustic interaction effects of installation, propulsion airframe aeroacoustics, will become more important as noise reduction targets become more difficult to achieve. In addition to continued fundamental component reduction efforts, a system level approach that includes propulsion airframe aeroacoustics will be required in order to achieve the 20 dB of perceived noise reduction envisioned by the long-range NASA goals. This emphasis on the aeroacoustics of propulsion airframe integration is a new part of NASA s noise research. The following paper will review current efforts and highlight technical challenges and approaches.

  19. Boeing 18-Inch Fan Rig Broadband Noise Test

    NASA Technical Reports Server (NTRS)

    Ganz, Ulrich W.; Joppa, Paul D.; Patten, Timothy J.; Scharpf, Daniel F.

    1998-01-01

    The purposes of the subject test were to identify and quantify the mechanisms by which fan broadband noise is produced, and to assess the validity of such theoretical models of those mechanisms as may be available. The test was conducted with the Boeing 18-inch fan rig in the Boeing Low-Speed Aeroacoustic Facility (LSAF). The rig was designed to be particularly clean and geometrically simple to facilitate theoretical modeling and to minimize sources of interfering noise. The inlet is cylindrical and is equipped with a boundary layer suction system. The fan is typical of modern high-by-pass ratio designs but is capable of operating with or without fan exit guide vanes (stators), and there is only a single flow stream. Fan loading and tip clearance are adjustable. Instrumentation included measurements of fan performance, the unsteady flow field incident on the fan and stators, and far-field and in-duct acoustic fields. The acoustic results were manipulated to estimate the noise generated by different sources. Significant fan broadband noise was found to come from the rotor self-noise as measured with clean inflow and no boundary layer. The rotor tip clearance affected rotor self-noise somewhat. The interaction of the rotor with inlet boundary layer turbulence is also a significant source, and is strongly affected by rotor tip clearance. High level noise can be generated by a high-order nonuniform rotating at a fraction of the fan speed, at least when tip clearance and loading are both large. Stator-generated noise is the loudest of the significant sources, by a small margin, at least on this rig. Stator noise is significantly affected by propagation through the fan.

  20. Helicopter tail rotor blade-vortex interaction noise

    NASA Technical Reports Server (NTRS)

    George, Albert R.; Chou, S.-T.

    1987-01-01

    A study is made of helicopter tail rotor noise, particularly that due to the interactions with main rotor tip vortices. Summarized here are present analysis, the computer codes, and the results of several test cases. Amiet's unsteady thin airfoil theory is used to calculate the acoustics of blade-vortex interaction. The noise source is modelled as a force dipole resulting from an airfoil of infinite span chopping through a skewed line vortex. To analyze the interactions between helicopter tail rotor and main rotor tip vortices, we developed a two-step approach: (1) the main rotor tip vortex system is obtained through a free wake geometry calculation of the main rotor using CAMRAD code; (2) acoustic analysis takes the results from the aerodynamic interaction analysis and calculates the farfield pressure signatures for the interactions. It is found that under a wide range of helicopter flight conditions, acoustic pressure fluctuations of significant magnitude can be generated by tail rotors due to a series of interactions with main rotor tip vortices. This noise mechanism depends strongly on the helicopter flight conditions and the relative location and phasing of the main and tail rotors. fluctuations of significant magnitude can be generated by tail rotors due to a series of interactions with main rotor tip vortices. This noise mechanism depends strongly upon the helicopter flight conditions and the relative location and phasing of the main and tail rotors.

  1. Flywheel rotor and containment technology development

    SciTech Connect

    Kulkarni, S.V.

    1981-08-11

    The goals of the project are: to develop an economical and practical composite flywheel having an energy density of 88 Wh/kg at failure, an operational energy density of 44 to 55 Wh/kg, and an energy storage capacity of approximately 1 kWh; to determine the suitability of various manufacturing processes for low-cost rotor fabrication; to investigate flywheel and flywheel-systems dynamics; to test and evaluate prototype rotors for use in transportation and stationary applications; and to develop a fail-safe, lightweight, and low-cost flywheel containment. The following tasks have been accomplished: evaluation and selection of 1-kWh, first-generation, advanced flywheel rotor designs for subsequent development towards the DOE-established energy density goal of 88 Wh/kg at burst; completion of an advanced design concept for a flywheel primary containment structure, capable of containing the failure of a 1-kWh flywheel rotor and targeted for vehicular applications; non-destructive inspection and burst testing of approximately twenty (20) prototype rotors, and initiation of cyclic testing; completion of various activities in the areas of rotor manufacturing processes, dynamic analyses and composite materials design data generation; and initiation of an economic feasibility study to establish a rational costing methodology for composite rotors and containment.

  2. High speed turboprop aeroacoustic study (counterrotation). Volume 1: Model development

    NASA Technical Reports Server (NTRS)

    Whitfield, C. E.; Mani, R.; Gliebe, P. R.

    1990-01-01

    The isolated counterrotating high speed turboprop noise prediction program was compared with model data taken in the GE Aircraft Engines Cell 41 anechoic facility, the Boeing Transonic Wind Tunnel, and in NASA-Lewis' 8x6 and 9x15 wind tunnels. The predictions show good agreement with measured data under both low and high speed simulated flight conditions. The installation effect model developed for single rotation, high speed turboprops was extended to include counterotation. The additional effect of mounting a pylon upstream of the forward rotor was included in the flow field modeling. A nontraditional mechanism concerning the acoustic radiation from a propeller at angle of attach was investigated. Predictions made using this approach show results that are in much closer agreement with measurement over a range of operating conditions than those obtained via traditional fluctuating force methods. The isolated rotors and installation effects models were combines into a single prediction program, results of which were compared with data taken during the flight test of the B727/UDF engine demonstrator aircraft. Satisfactory comparisons between prediction and measured data for the demonstrator airplane, together with the identification of a nontraditional radiation mechanism for propellers at angle of attack are achieved.

  3. Aeroacoustics research in Europe: The CEAS-ASC report on 2013 highlights

    NASA Astrophysics Data System (ADS)

    Bennett, G. J.; Kennedy, J.; Meskell, C.; Carley, M.; Jordan, P.; Rice, H.

    2015-03-01

    The Council of European Aerospace Societies (CEAS) Aeroacoustics Specialists Committee (ASC) supports and promotes the interests of the scientific and industrial aeroacoustics community on an European scale and European aeronautics activities internationally. In this context, "aeroacoustics" encompasses all aerospace acoustics and related areas. Each year the committee highlights some of the research and development projects in Europe. This paper is a report on highlights of aeroacoustics research in Europe in 2013, compiled from information provided to the ASC of the CEAS. During 2013, a number of research programmes involving aeroacoustics were funded by the European Commission. Some of the highlights from these programmes are summarised in this paper, as well as highlights from other programmes funded by national programmes or by industry. Furthermore, a concise summary of the CEAS-ASC workshop "Atmospheric and Ground Effects on Aircraft Noise" held in Seville, Spain in September 2013 is included in this report. Enquiries concerning all contributions should be addressed to the authors who are given at the end of each subsection. This issue of the "highlights" paper is dedicated to the memory of Prof. John A. Fitzpatrick, Professor of Mechanical Engineering, Trinity College Dublin, and a valued member of the Aeroacoustics Specialists Committee. John passed away in September 2012 and is fondly missed across the globe by the friends he made in the Aeroacoustics Community. This paper is edited by PhD graduates and colleagues of John's who conduct research in aeroacoustics, inspired by his thirst for knowledge.

  4. On the flow field around a Savonius rotor

    NASA Astrophysics Data System (ADS)

    Bergeles, G.; Athanassiadis, N.

    A model of a two-bucket Savonius rotor windmill was constructed and tested in a wind tunnel. The flow field around the rotor was examined visually and also quantitatively with the use of a hot wire. The flow visualization revealed an upstream influence on the flow field up to 3 rotor diameters away and a strong downwash downstream. Hot wire measurements showed a large velocity deficit behind the rotor and a quick velocity recovery downstream due to strong mixing; the latter was associated with high levels of turbulence. Energy spectra revealed that all turbulence was concentrated in a single harmonic corresponding to twice the rotational speed of the rotor.

  5. Aeroacoustic Study of a High-Fidelity Aircraft Model: Part 1- Steady Aerodynamic Measurements

    NASA Technical Reports Server (NTRS)

    Khorrami, Mehdi R.; Hannon, Judith A.; Neuhart, Danny H.; Markowski, Gregory A.; VandeVen, Thomas

    2012-01-01

    In this paper, we present steady aerodynamic measurements for an 18% scale model of a Gulfstream air-craft. The high fidelity and highly-instrumented semi-span model was developed to perform detailed aeroacoustic studies of airframe noise associated with main landing gear/flap components and gear-flap interaction noise, as well as to evaluate novel noise reduction concepts. The aeroacoustic tests, being conducted in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel, are split into two entries. The first entry, completed November 2010, was entirely devoted to the detailed mapping of the aerodynamic characteristics of the fabricated model. Flap deflections of 39?, 20?, and 0? with the main landing gear on and off were tested at Mach numbers of 0.16, 0.20, and 0.24. Additionally, for each flap deflection, the model was tested with the tunnel both in the closed-wall and open-wall (jet) modes. During this first entry, global forces (lift and drag) and extensive steady and unsteady surface pressure measurements were obtained. Preliminary analysis of the measured forces indicates that lift, drag, and stall characteristics compare favorably with Gulfstream?s high Reynolds number flight data. The favorable comparison between wind-tunnel and flight data allows the semi-span model to be used as a test bed for developing/evaluating airframe noise reduction concepts under a relevant environment. Moreover, initial comparison of the aerodynamic measurements obtained with the tunnel in the closed- and open-wall configurations shows similar aerodynamic behavior. This permits the acoustic and off-surface flow measurements, planned for the second entry, to be conducted with the tunnel in the open-jet mode.

  6. Assessment of Hybrid RANS/LES Turbulence Models for Aeroacoustics Applications

    NASA Technical Reports Server (NTRS)

    Vatsa, Veer N.; Lockhard, David P.

    2010-01-01

    Predicting the noise from aircraft with exposed landing gear remains a challenging problem for the aeroacoustics community. Although computational fluid dynamics (CFD) has shown promise as a technique that could produce high-fidelity flow solutions, generating grids that can resolve the pertinent physics around complex configurations can be very challenging. Structured grids are often impractical for such configurations. Unstructured grids offer a path forward for simulating complex configurations. However, few unstructured grid codes have been thoroughly tested for unsteady flow problems in the manner needed for aeroacoustic prediction. A widely used unstructured grid code, FUN3D, is examined for resolving the near field in unsteady flow problems. Although the ultimate goal is to compute the flow around complex geometries such as the landing gear, simpler problems that include some of the relevant physics, and are easily amenable to the structured grid approaches are used for testing the unstructured grid approach. The test cases chosen for this study correspond to the experimental work on single and tandem cylinders conducted in the Basic Aerodynamic Research Tunnel (BART) and the Quiet Flow Facility (QFF) at NASA Langley Research Center. These configurations offer an excellent opportunity to assess the performance of hybrid RANS/LES turbulence models that transition from RANS in unresolved regions near solid bodies to LES in the outer flow field. Several of these models have been implemented and tested in both structured and unstructured grid codes to evaluate their dependence on the solver and mesh type. Comparison of FUN3D solutions with experimental data and numerical solutions from a structured grid flow solver are found to be encouraging.

  7. Experimental hingeless rotor characteristics at low advance ratio with thrust. [wind tunnel tests of rotary wing operating at moderate to high lift

    NASA Technical Reports Server (NTRS)

    London, R. J.; Watts, G. A.; Sissingh, G. J.

    1973-01-01

    An experimental investigation to determine the dynamic characteristics of a hingeless rotor operating at moderate to high lift was conducted on a small scale, 7.5-foot diameter, four-bladed hingeless rotor model in a 7 x 10-foot wind tunnel. The primary objective of this research program was the empirical determination of the rotor steady-state and frequency responses to swashplate and body excitations. Collective pitch was set from 0 to 20 degrees, with the setting at a particular advance ratio limited by the cyclic pitch available for hub moment trim. Advance ratio varied from 0.00 to 0.36 for blades with nondimensional first-flap frequencies at 1.15, 1.28 and 1.33 times the rotor rotation frequency. Several conditions were run with the rotor operating in the transition regime. Rotor response at high lift is shown to be generally nonlinear in this region. As a secondary objective an experimental investigation of the rotor response to 4/revolution swashplate excitations at advance ratios of 0.2 to 0.85 and at a nondimensional, first-flap modal frequency of 1.34 was also conducted, using the 7 x 10-foot wind tunnel. It is shown that 4/revolution swashplate inputs are a method for substantially reducing rotor-induced, shafttransmitted vibratory forces.

  8. Experimental study of main rotor tip geometry and tail rotor interactions in hover. Volume 1. Text and figures

    NASA Technical Reports Server (NTRS)

    Balch, D. T.; Lombardi, J.

    1985-01-01

    A model scale hover test was conducted in the Sikorsky Aircraft Model rotor hover Facility to identify and quantify the impact of the tail rotor on the demonstrated advantages of advanced geometry tip configurations. The test was conducted using the Basic Model Test Rig and two scaled main rotor systems, one representing a 1/5.727 scale UH-60A BLACK HAWK and the others a 1/4.71 scale S-76. Eight alternate rotor tip configurations were tested, 3 on the BLACK HAWK rotor and 6 on the S-76 rotor. Four of these tips were then selected for testing in close proximity to an operating tail rotor (operating in both tractor and pusher modes) to determine if the performance advantages that could be obtained from the use of advanced geometry tips in a main rotor only environment would still exist in the more complex flow field involving a tail rotor. The test showed that overall the tail rotor effects on the advanced tip configurations tested are not substantially different from the effects on conventional tips.

  9. Modelling Aerodynamically Generated Sound: Recent Advances in Rotor Noise Prediction

    NASA Technical Reports Server (NTRS)

    Brentner, Kenneth S.

    2000-01-01

    A great deal of progress has been made in the modeling of aerodynamically generated sound for rotors over the past decade. The Ffowcs Williams-Hawkings (FW-H ) equation has been the foundation for much of the development. Both subsonic and supersonic quadrupole noise formulations have been developed for the prediction of high-speed impulsive noise. In an effort to eliminate the need to compute the quadrupole contribution, the FW-H has also been utilized on permeable surfaces surrounding all physical noise sources. Comparison of the Kirchhoff formulation for moving surfaces with the FW-H equation have shown that the Kirchhoff formulation for moving surfaces can give erroneous results for aeroacoustic problems.

  10. Aeroacoustics research in Europe: The CEAS-ASC report on 2012 highlights

    NASA Astrophysics Data System (ADS)

    Bodén, H.; Efraimsson, G.

    2013-12-01

    The Council of European Aerospace Societies (CEAS) Aeroacoustics Specialists Committee (ASC) supports and promotes the interests of the scientific and industrial aeroacoustics community on an European scale and European aeronautics activities internationally. In this context, "aeroacoustics" encompasses all aerospace acoustics and related areas. Each year the committee highlights some of the research and development projects in Europe. This paper is a report on highlights of aeroacoustics research in Europe in 2012, compiled from information provided to the ASC of the CEAS. During 2012, a number of research programmes involving aeroacoustics were funded by the European Commission. Some of the highlights from these programmes are summarized in this paper, as well as highlights from other programmes funded by national programmes or by industry. Enquiries concerning all contributions should be addressed to the authors who are given at the end of each subsection.

  11. Fourth Computational Aeroacoustics (CAA) Workshop on Benchmark Problems

    NASA Technical Reports Server (NTRS)

    Dahl, Milo D. (Editor)

    2004-01-01

    This publication contains the proceedings of the Fourth Computational Aeroacoustics (CAA) Workshop on Benchmark Problems. In this workshop, as in previous workshops, the problems were devised to gauge the technological advancement of computational techniques to calculate all aspects of sound generation and propagation in air directly from the fundamental governing equations. A variety of benchmark problems have been previously solved ranging from simple geometries with idealized acoustic conditions to test the accuracy and effectiveness of computational algorithms and numerical boundary conditions; to sound radiation from a duct; to gust interaction with a cascade of airfoils; to the sound generated by a separating, turbulent viscous flow. By solving these and similar problems, workshop participants have shown the technical progress from the basic challenges to accurate CAA calculations to the solution of CAA problems of increasing complexity and difficulty. The fourth CAA workshop emphasized the application of CAA methods to the solution of realistic problems. The workshop was held at the Ohio Aerospace Institute in Cleveland, Ohio, on October 20 to 22, 2003. At that time, workshop participants presented their solutions to problems in one or more of five categories. Their solutions are presented in this proceedings along with the comparisons of their solutions to the benchmark solutions or experimental data. The five categories for the benchmark problems were as follows: Category 1:Basic Methods. The numerical computation of sound is affected by, among other issues, the choice of grid used and by the boundary conditions. Category 2:Complex Geometry. The ability to compute the sound in the presence of complex geometric surfaces is important in practical applications of CAA. Category 3:Sound Generation by Interacting With a Gust. The practical application of CAA for computing noise generated by turbomachinery involves the modeling of the noise source mechanism as a

  12. Aero-acoustic performance comparison of core engine noise suppressors on NASA quiet engine C

    NASA Technical Reports Server (NTRS)

    Bloomer, H. E.; Schaefer, J. W.

    1977-01-01

    The relative aero-acoustic effectiveness of two core engine suppressors, a contractor-designed suppressor delivered with the Quiet Engine, and a NASA-designed suppressor was evaluated. The NASA suppressor was tested with and without a splitter making a total of three configurations being reported in addition to the baseline hardwall case. The aerodynamic results are presented in terms of tailpipe pressure loss, corrected net thrust, and corrected specific fuel consumption as functions of engine power setting. The acoustic results are divided into duct and far-field acoustic data. The NASA-designed core suppressor did the better job of suppressing aft end noise, but the splitter associated with it caused a significant engine performance penality. The NASA core suppressor without the spltter suppressed most of the core noise without any engine performance penalty.

  13. The boundary data immersion method for compressible flows with application to aeroacoustics

    NASA Astrophysics Data System (ADS)

    Schlanderer, Stefan C.; Weymouth, Gabriel D.; Sandberg, Richard D.

    2017-03-01

    This paper introduces a virtual boundary method for compressible viscous fluid flow that is capable of accurately representing moving bodies in flow and aeroacoustic simulations. The method is the compressible extension of the boundary data immersion method (BDIM, Maertens & Weymouth (2015), [18]). The BDIM equations for the compressible Navier-Stokes equations are derived and the accuracy of the method for the hydrodynamic representation of solid bodies is demonstrated with challenging test cases, including a fully turbulent boundary layer flow and a supersonic instability wave. In addition we show that the compressible BDIM is able to accurately represent noise radiation from moving bodies and flow induced noise generation without any penalty in allowable time step.

  14. Aero-acoustic performance comparison of core engine noise suppressors on NASA quiet engine 'C'

    NASA Technical Reports Server (NTRS)

    Bloomer, H. E.; Schaefer, J. W.

    1977-01-01

    The purpose of the experimental program reported herein was to evaluate and compare the relative aero-acoustic effectiveness of two core engine suppressors, a contractor-designed suppressor delivered with the Quiet Engine, and a NASA-designed suppressor, designed and built subsequently. The NASA suppressor was tested with and without a splitter making a total of three configurations being reported in addition to the baseline hardwall case. The aerodynamic results are presented in terms of tailpipe pressure loss, corrected net thrust, and corrected specific fuel consumption as functions of engine power setting. The acoustic results are divided into duct and far-field acoustic data. The NASA-designed core suppressor did the better job of suppressing aft end noise, but the splitter associated with it caused a significant engine performance penalty. The NASA core suppressor without the splitter suppressed most of the core noise without any engine performance penalty.

  15. Aeroacoustic research programs at the Army Aviation Research and Technology Activity

    NASA Technical Reports Server (NTRS)

    Yu, Yung H.; Schmitz, Fredric H.; Morse, H. Andrew

    1988-01-01

    The Army rotorcraft aeroacoustic programs are reviewed, highlighting the theoretical and experimental progress made by Army researchers in the physical understanding of helicopter impulsive noise. The two impulsive noise sources addressed over this past decade are high-speed impulsive noise and blade-vortex interaction noise, both of which have had and will continue to have an increasing influence on Army rotorcraft design and operations. The advancements discussed are in the areas of in-flight data acquisition techniques, small-scale-model tests in wind tunnels, holographic interferometry/tomographic techniques, and the expanding capabilities of computational fluid dynamics in rotorcraft acoustic problems. Current theoretical prediction methods are compared with experimental data, and parameters that govern model scaling are established. The very successful cooperative efforts between the Army, NASA, and industry are also addressed

  16. Investigation of computational and spectral analysis methods for aeroacoustic wave propagation

    NASA Technical Reports Server (NTRS)

    Vanel, Florence O.

    1995-01-01

    Most computational fluid dynamics (CFD) schemes are not adequately accurate for solving aeroacoustics problems, which have wave amplitudes several orders of magnitude smaller yet with frequencies larger than the flow field variations generating the sound. Hence, a computational aeroacoustics (CAA) algorithm should have minimal dispersion and dissipation features. A dispersion relation preserving (DRP) scheme is, therefore, applied to solve the linearized Euler equations in order to simulate the propagation of three types of waves, namely: acoustic, vorticity, and entropy waves. The scheme is derived using an optimization procedure to ensure that the numerical derivatives preserve the wave number and angular frequency of the partial differential equations being discretized. Consequently, simulated waves propagate with the correct wave speeds and exhibit their appropriate properties. A set of radiation and outflow boundary conditions, compatible with the DRP scheme and derived from the asymptotic solutions of the governing equations, are also implemented. Numerical simulations are performed to test the effectiveness of the DRP scheme and its boundary conditions. The computed solutions are shown to agree favorably with the exact solutions. The major restriction appears to be that the dispersion relations can be preserved only for waves with wave lengths longer than four or five spacings. The boundary conditions are found to be transparent to the outgoing disturbances. However, when the disturbance source is placed closer to a boundary, small acoustic reflections start appearing. CAA generates enormous amounts of temporal data which needs to be reduced to understand the physical problem being simulated. Spectral analysis is one approach that helps us in extracting information which often can not be easily interpreted in the time domain. Thus, three different methods for the spectral analysis of numerically generated aeroacoustic data are studied. First, the

  17. Rotor control system

    NASA Technical Reports Server (NTRS)

    Bradford, Michael P. (Inventor); Maciolek, Joseph R. (Inventor)

    1987-01-01

    A helicopter rotor control system (13) including a stop azimuth controller (32) for establishing the value of a deceleration command (15') to a deceleration controller (23), a transition azimuth predictor (41) and a position reference generator (55), which are effective during the last revolution of said rotor (14) to establish a correction indication (38) to adjust the deceleration command (15') to ensure that one of the rotor blades (27) stops at a predetermined angular position.

  18. Effects of ingested atmospheric turbulence on measured tail rotor acoustics

    NASA Technical Reports Server (NTRS)

    Signor, David B.; Yamauchi, Gloria K.; Mosher, Marianne; Hagen, Martin J.; George, Albert R.

    1992-01-01

    Results from an outdoor hover test of a full-scale Lynx tail rotor are presented. The investigation was designed to further the understanding of the acoustics of an isolated tail rotor hovering out-of-ground effect in atmospheric turbulence, without the effects of the main rotor wake or other helicopter components. Measurements include simultaneous rotor performance, noise, inflow, and far-field atmospheric turbulence. Results with grid-generated inflow turbulence are also presented. The effects of turbulence ingestion on rotor noise are quantified. Turbulence ingestion noise is found to be the dominant noise mechanism at locations near the rotor axis. At these locations, the sound radiated by the hovering rotor increases with both increasing atmospheric wind speed and ingested rms turbulent velocity.

  19. Aeroacoustic wind-tunnel tests of a light twin-boom general-aviation airplane with free or shrouded-pusher propellers. [in the Langley full-scale tunnel

    NASA Technical Reports Server (NTRS)

    Mclemore, H. C.; Pegg, R. J.

    1980-01-01

    Tests were conducted in the Langley full-scale tunnel to determine the aerodynamic performance and acoustic characteristics of four different pusher-propeller configurations on a twin boom, general aviation airplane. The propellers included a 2-blade free propeller, two 3-blade shrouded propellers, and a 5-blade shrouded propeller. The tests were conducted for a range of airplane angles of attack from about 0 deg to 16 deg for test speeds from 0 to about 36 m/sec and for a range of propeller blade angles and rotation speeds. The free propeller provided the best aerodynamic propulsive performance. For forward flight conditions, the free propeller noise levels were lower than those of the shrouded propellers. In the static conditions the free propeller noise levels were as low as those for the shrouded propellers, except for the propeller in-plane noise where the shrouded propeller noise levels were lower.

  20. Aero-acoustic experimental verification of optimum configuration of variable-pitch fans for 40 x 80 foot subsonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Lown, H.

    1977-01-01

    The aerodynamic and acoustic performance of two drive fan configurations (low-speed and high-speed variable pitch design) for a 40 x 80 foot wind tunnel were monitored. A 1/7-scale model was utilized. The necessary aero-acoustic data reduction computer program logic was supplied. Test results were evaluated, and the optimum configuration to be employed in the 40 foot full scale fan was recommended.

  1. Advanced Model for Extreme Lift and Improved Aeroacoustics (AMELIA)

    NASA Technical Reports Server (NTRS)

    Lichtwardt, Jonathan; Paciano, Eric; Jameson, Tina; Fong, Robert; Marshall, David

    2012-01-01

    tunnel model design would be completed, manufactured, and calibrated. During the fifth year the large scale wind tunnel test was conducted. This technical memo will describe all phases of the Advanced Model for Extreme Lift and Improved Aeroacoustics (AMELIA) project and provide a brief summary of the background and modeling efforts involved in the NRA. The conceptual designs considered for this project and the decision process for the selected configuration adapted for a wind tunnel model will be briefly discussed. The internal configuration of AMELIA, and the internal measurements chosen in order to satisfy the requirements of obtaining a database of experimental data to be used for future computational model validations. The external experimental techniques that were employed during the test, along with the large-scale wind tunnel test facility are covered in great detail. Experimental measurements in the database include forces and moments, and surface pressure distributions, local skin friction measurements, boundary and shear layer velocity profiles, far-field acoustic data and noise signatures from turbofan propulsion simulators. Results and discussion of the circulation control performance, over-the-wing mounted engines, and the combined performance are also discussed in great detail.

  2. Experimental study of main rotor/tail rotor/airframe interactions in hover. Volume 1: Text and figures

    NASA Technical Reports Server (NTRS)

    Balch, D. T.; Saccullo, A.; Sheehy, T. W.

    1983-01-01

    To assist in identifying and quantifying the relevant parameters associated with the complex topic of main rotor/fuselage/tail rotor interference, a model scale hover test was conducted in the Model Rotor Hover Facility. The test was conducted using the basic model test rig, fuselage skins to represent a UH-60A BLACK HAWK helicopter, 4 sets of rotor blades of varying geometry (i.e., twist, airfoils and solidity) and a model tail rotor that could be relocated to give changes in rotor clearance (axially, laterally, and vertically), can't angle and operating model (pusher or tractor). The description of the models and the tests, data analysis and summary (including plots) are included. The customary system of units gas used for principal measurements and calculations. Expressions in both SI units and customary units are used with the SI units stated first and the customary units afterwords, in parenthesis.

  3. Benchmark Problems Used to Assess Computational Aeroacoustics Codes

    NASA Technical Reports Server (NTRS)

    Dahl, Milo D.; Envia, Edmane

    2005-01-01

    The field of computational aeroacoustics (CAA) encompasses numerical techniques for calculating all aspects of sound generation and propagation in air directly from fundamental governing equations. Aeroacoustic problems typically involve flow-generated noise, with and without the presence of a solid surface, and the propagation of the sound to a receiver far away from the noise source. It is a challenge to obtain accurate numerical solutions to these problems. The NASA Glenn Research Center has been at the forefront in developing and promoting the development of CAA techniques and methodologies for computing the noise generated by aircraft propulsion systems. To assess the technological advancement of CAA, Glenn, in cooperation with the Ohio Aerospace Institute and the AeroAcoustics Research Consortium, organized and hosted the Fourth CAA Workshop on Benchmark Problems. Participants from industry and academia from both the United States and abroad joined to present and discuss solutions to benchmark problems. These demonstrated technical progress ranging from the basic challenges to accurate CAA calculations to the solution of CAA problems of increasing complexity and difficulty. The results are documented in the proceedings of the workshop. Problems were solved in five categories. In three of the five categories, exact solutions were available for comparison with CAA results. A fourth category of problems representing sound generation from either a single airfoil or a blade row interacting with a gust (i.e., problems relevant to fan noise) had approximate analytical or completely numerical solutions. The fifth category of problems involved sound generation in a viscous flow. In this case, the CAA results were compared with experimental data.

  4. NASA's Aeroacoustic Tools and Methods for Analysis of Aircraft Noise

    NASA Technical Reports Server (NTRS)

    Rizzi, Stephen A.; Lopes, Leonard V.; Burley, Casey L.

    2015-01-01

    Aircraft community noise is a significant concern due to continued growth in air traffic, increasingly stringent environmental goals, and operational limitations imposed by airport authorities. The ability to quantify aircraft noise at the source and ultimately at observers is required to develop low noise aircraft designs and flight procedures. Predicting noise at the source, accounting for scattering and propagation through the atmosphere to the observer, and assessing the perception and impact on a community requires physics-based aeroacoustics tools. Along with the analyses for aero-performance, weights and fuel burn, these tools can provide the acoustic component for aircraft MDAO (Multidisciplinary Design Analysis and Optimization). Over the last decade significant progress has been made in advancing the aeroacoustic tools such that acoustic analyses can now be performed during the design process. One major and enabling advance has been the development of the system noise framework known as Aircraft NOise Prediction Program2 (ANOPP2). ANOPP2 is NASA's aeroacoustic toolset and is designed to facilitate the combination of acoustic approaches of varying fidelity for the analysis of noise from conventional and unconventional aircraft. The toolset includes a framework that integrates noise prediction and propagation methods into a unified system for use within general aircraft analysis software. This includes acoustic analyses, signal processing and interfaces that allow for the assessment of perception of noise on a community. ANOPP2's capability to incorporate medium fidelity shielding predictions and wind tunnel experiments into a design environment is presented. An assessment of noise from a conventional and Hybrid Wing Body (HWB) aircraft using medium fidelity scattering methods combined with noise measurements from a model-scale HWB recently placed in NASA's 14x22 wind tunnel are presented. The results are in the form of community noise metrics and

  5. Application of cylindrical near-field acoustical holography to the visualization of aeroacoustic sources.

    PubMed

    Lee, Moohyung; Bolton, J Stuart; Mongeau, Luc

    2003-08-01

    The purpose of this study was to develop methods for visualizing the sound radiation from aeroacoustic sources in order to identify their source strength distribution, radiation patterns, and to quantify the performance of noise control solutions. Here, cylindrical Near-field Acoustical Holography was used for that purpose. In a practical holographic measurement of sources comprising either partially correlated or uncorrelated subsources, it is necessary to use a number of reference microphones so that the sound field on the hologram surface can be decomposed into mutually incoherent partial fields before holographic projection. In this article, procedures are described for determining the number of reference microphones required when visualizing partially correlated aeroacoustic sources; performing source nonstationarity compensation; and applying regularization. The procedures have been demonstrated by application to a ducted fan. Holographic tests were performed to visualize the sound radiation from that source in its original form. The system was then altered to investigate the effect of two modifications on the fan's sound radiation pattern: first, leaks were created in the fan and duct assembly, and second, sound absorbing material was used to line the downstream duct section. Results in all three cases are shown at the blade passing frequency and for a broadband noise component. In the absence of leakage, both components were found to exhibit a dipole-like radiation pattern. Leakage was found to have a strong influence on the directivity of the blade passing tone. The increase of the flow resistance caused by adding the acoustical lining resulted in a nearly symmetric reduction of sound radiation.

  6. Identifying equivalent sound sources from aeroacoustic simulations using a numerical phased array

    NASA Astrophysics Data System (ADS)

    Pignier, Nicolas J.; O'Reilly, Ciarán J.; Boij, Susann

    2017-04-01

    An application of phased array methods to numerical data is presented, aimed at identifying equivalent flow sound sources from aeroacoustic simulations. Based on phased array data extracted from compressible flow simulations, sound source strengths are computed on a set of points in the source region using phased array techniques assuming monopole propagation. Two phased array techniques are used to compute the source strengths: an approach using a Moore-Penrose pseudo-inverse and a beamforming approach using dual linear programming (dual-LP) deconvolution. The first approach gives a model of correlated sources for the acoustic field generated from the flow expressed in a matrix of cross- and auto-power spectral values, whereas the second approach results in a model of uncorrelated sources expressed in a vector of auto-power spectral values. The accuracy of the equivalent source model is estimated by computing the acoustic spectrum at a far-field observer. The approach is tested first on an analytical case with known point sources. It is then applied to the example of the flow around a submerged air inlet. The far-field spectra obtained from the source models for two different flow conditions are in good agreement with the spectra obtained with a Ffowcs Williams-Hawkings integral, showing the accuracy of the source model from the observer's standpoint. Various configurations for the phased array and for the sources are used. The dual-LP beamforming approach shows better robustness to changes in the number of probes and sources than the pseudo-inverse approach. The good results obtained with this simulation case demonstrate the potential of the phased array approach as a modelling tool for aeroacoustic simulations.

  7. Performance and Vibratory Loads Data From a Wind-Tunnel Test of a Model Helicopter Main-Rotor Blade With a Paddle-Type Tip

    NASA Technical Reports Server (NTRS)

    Yeager, William T., Jr.; Noonan, Kevin W.; Singleton, Jeffrey D.; Wilbur, Matthew L.; Mirick, Paul H.

    1997-01-01

    An investigation was conducted in the Langley Transonic Dynamics Tunnel to obtain data to permit evaluation of paddle-type tip technology for possible use in future U.S. advanced rotor designs. Data was obtained for both a baseline main-rotor blade and a main-rotor blade with a paddle-type tip. The baseline and paddle-type tip blades were compared with regard to rotor performance, oscillatory pitch-link loads, and 4-per-rev vertical fixed-system loads. Data was obtained in hover and forward flight over a nominal range of advance ratios from 0.15 to 0.425. Results indicate that the paddle-type tip offers no performance improvements in either hover or forward flight. Pitch-link oscillatory loads for the paddle-type tip are higher than for the baseline blade, whereas 4-per-rev vertical fixed-system loads are generally lower.

  8. An assessment of artificial damping models for aeroacoustic calculations

    NASA Technical Reports Server (NTRS)

    Hayder, M. Ehtesham

    1995-01-01

    We present a study of the effect of artificial dissipation models on nonlinear wave computations using a few high order schemes. Our motivation is to assess the effectiveness of artificial dissipation models for their suitability for aeroacoustic computations. We solve three model problems in one dimension using the Euler equations. Initial conditions are chosen to generate nonlinear waves in the computational domain. We examine various dissipation models in central difference schemes such as the Dispersion Relation Preserving (DRP) scheme and the standard fourth and sixth order schemes. We also make a similar study with the fourth order MacCormack scheme due to Gottieb and Turkel.

  9. High-intensity tone generation by aeroacoustic sources

    NASA Technical Reports Server (NTRS)

    Shakkottai, P.; Kwack, E. Y.; Cho, Y. I.; Back, L. H.

    1987-01-01

    An experimental investigation has been carried out on the production of high-intensity tones by axisymmetric ring cavities. Maximum sound production occurs during an acoustic resonance at Strouhal numbers, which depend only on the local flow velocity independent of cavity location. Values of sound pressure of about 115 dB at 1-m distance can be generated by axisymmetric ring cavities on projectiles moving at a relatively low flight speed equal to 70 m/s. Frequencies in the audible range up to several kilohertz can be generated aeroacoustically. A simple analytical model has been developed to explain the experimental observations.

  10. Program in acoustics. [aeroacoustics, aircraft noise, and noise suppression

    NASA Technical Reports Server (NTRS)

    1979-01-01

    Relevant research projects conducted by faculty and graduate students in the general area of aeroacoustics to further the understanding of noise generation by aircraft and to aid in the development of practical methods for noise suppression are listed. Special activities summarized relate to the nonlinear acoustic wave theory and its application to several cases including that of the acoustic source located at the throat of a near-sonic duct, a computer program developed to compute the nonlinear wave theory, and a parabolic approximation for propagation of sounding in moving stratified media.

  11. Time-Domain Impedance Boundary Conditions for Computational Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Tam, Christopher K. W.; Auriault, Laurent

    1996-01-01

    It is an accepted practice in aeroacoustics to characterize the properties of an acoustically treated surface by a quantity known as impedance. Impedance is a complex quantity. As such, it is designed primarily for frequency-domain analysis. Time-domain boundary conditions that are the equivalent of the frequency-domain impedance boundary condition are proposed. Both single frequency and model broadband time-domain impedance boundary conditions are provided. It is shown that the proposed boundary conditions, together with the linearized Euler equations, form well-posed initial boundary value problems. Unlike ill-posed problems, they are free from spurious instabilities that would render time-marching computational solutions impossible.

  12. Towards Numerical Simulations of Trailing-Edge Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Wang, Meng

    1996-01-01

    The aeroacoustics of flow-hydrofoil interactions exhibits distinct characteristics depending on the physical length scales involved. In the small-foil (relative to acoustic wavelength) limit characteristic of the noise generated by large-scale vortex shedding at low flow Mach number, the noise calculation is facilitated by the use of the Lighthill analogy in conjunction with a free-space Green's function, in the sense of Curle's formulation. A methodology for computing the vortex-shedding noise using the Curle formulation, including both surface-induced dipole sources and volume quadrupole sources, has been developed.

  13. Application of computational aero-acoustics to real world problems

    NASA Technical Reports Server (NTRS)

    Hardin, Jay C.

    1996-01-01

    The application of computational aeroacoustics (CAA) to real problems is discussed in relation to the analysis performed with the aim of assessing the application of the various techniques. It is considered that the applications are limited by the inability of the computational resources to resolve the large range of scales involved in high Reynolds number flows. Possible simplifications are discussed. It is considered that problems remain to be solved in relation to the efficient use of the power of parallel computers and in the development of turbulent modeling schemes. The goal of CAA is stated as being the implementation of acoustic design studies on a computer terminal with reasonable run times.

  14. Structural qualification testing and operational loading on a fiberglass rotor blade for the Mod-OA wind turbine

    NASA Technical Reports Server (NTRS)

    Sullivan, T. L.

    1983-01-01

    Fatigue tests were performed on full- and half-scale root end sections, first to qualify the root retention design, and second to induce failure. Test methodology and results are presented. Two operational blades were proof tested to design limit load to ascertain buckling resistance. Measurements of natural frequency, damping ratio, and deflection under load made on the operational blades are documented. The tests showed that all structural design requirements were met or exceeded. Blade loads measured during 3000 hr of field operation were close to those expected. The measured loads validated the loads used in the fatigue tests and gave high confidence in the ability of the blades to achieve design life.

  15. Full-scale hingeless rotor performance and loads

    NASA Technical Reports Server (NTRS)

    Peterson, Randall L.

    1995-01-01

    A full-scale BO-105 hingeless rotor system was tested in the NASA Ames 40- by 80-Foot Wind Tunnel on the rotor test apparatus. Rotor performance, rotor loads, and aeroelastic stability as functions of both collective and cyclic pitch, tunnel velocity, and shaft angle were investigated. This test was performed in support of the Rotor Data Correlation Task under the U.S. Army/German Memorandum of Understanding on Cooperative Research in the Field of Helicopter Aeromechanics. The primary objective of this test program was to create a data base for full-scale hingeless rotor performance and structural blade loads. A secondary objective was to investigate the ability to match flight test conditions in the wind tunnel. This data base can be used for the experimental and analytical studies of hingeless rotor systems over large variations in rotor thrust and tunnel velocity. Rotor performance and structural loads for tunnel velocities from hover to 170 knots and thrust coefficients (C(sub T)/sigma) from 0.0 to 0.12 are presented in this report. Thrust sweeps at tunnel velocities of 10, 20, and 30 knots are also included in this data set.

  16. Estimation of dynamic rotor loads for the rotor systems research aircraft: Methodology development and validation

    NASA Technical Reports Server (NTRS)

    Duval, R. W.; Bahrami, M.

    1985-01-01

    The Rotor Systems Research Aircraft uses load cells to isolate the rotor/transmission systm from the fuselage. A mathematical model relating applied rotor loads and inertial loads of the rotor/transmission system to the load cell response is required to allow the load cells to be used to estimate rotor loads from flight data. Such a model is derived analytically by applying a force and moment balance to the isolated rotor/transmission system. The model is tested by comparing its estimated values of applied rotor loads with measured values obtained from a ground based shake test. Discrepancies in the comparison are used to isolate sources of unmodeled external loads. Once the structure of the mathematical model has been validated by comparison with experimental data, the parameters must be identified. Since the parameters may vary with flight condition it is desirable to identify the parameters directly from the flight data. A Maximum Likelihood identification algorithm is derived for this purpose and tested using a computer simulation of load cell data. The identification is found to converge within 10 samples. The rapid convergence facilitates tracking of time varying parameters of the load cell model in flight.

  17. NASA/HAA Advanced Rotorcraft Technology and Tilt Rotor Workshop. Volume 7: Tilt Rotor Session

    NASA Technical Reports Server (NTRS)

    1980-01-01

    The technical characteristics of the XV-15 aircraft were discussed. Program objectives, concept evaluation, tilt rotor experiments and civil market applications are presented. The XV-15 status and test schedule are also included.

  18. Development and testing of a 2.5 kW synchronous generator with a high temperature superconducting stator and permanent magnet rotor

    NASA Astrophysics Data System (ADS)

    Qu, Timing; Song, Peng; Yu, Xiaoyu; Gu, Chen; Li, Longnian; Li, Xiaohang; Wang, Dewen; Hu, Boping; Chen, Duxing; Zeng, Pan; Han, Zhenghe

    2014-04-01

    High temperature superconducting (HTS) armature windings have the potential for increasing the electric loading of a synchronous generator due to their high current transport capacity, which could increase the power density of an HTS rotating machine. In this work, a novel synchronous generator prototype with an HTS stator and permanent magnet rotor has been developed. It has a basic structure of four poles and six slots. The armature winding was constructed from six double-pancake race-track coils with 44 turns each. It was designed to deliver 2.5 kW at 300 rpm. A concentrated winding configuration was proposed, to prevent interference at the ends of adjacent HTS coils. The HTS stator was pressure mounted into a hollow Dewar cooled with liquid nitrogen. The whole stator could be cooled down to around 82 K by conduction cooling. In the preliminary testing, the machine worked properly and could deliver 1.8 kW power when the armature current was 14.4 A. Ic for the HTS coils was found to be suppressed due to the influence of the temperature and the leakage field.

  19. Influence of rubbing on rotor dynamics, part 2

    NASA Technical Reports Server (NTRS)

    Muszynska, Agnes; Bently, Donald E.; Franklin, Wesley D.; Hayashida, Robert D.; Kingsley, Lori M.; Curry, Arthur E.

    1989-01-01

    Rotor dynamic behavior depends considerably on how much the specific physical phenomena accompanying rotor rubbing against the stator is involved. The experimental results of rotor-to-stator rubbing contact are analyzed. The computer code is described for obtaining numerical calculations of rotor-to-stator rubbing system dynamic responses. Computer generated results are provided. The reduced dynamic data from High Pressure Fuel Turbo Pump (HPFTP) hot fire test are given. The results provide some significant conclusions. Information is provided on the electronic instrumentation used in the experimental testing.

  20. Instability thresholds for flexible rotors in hydrodynamic bearings

    NASA Technical Reports Server (NTRS)

    Allaire, P. E.; Flack, R. D.

    1980-01-01

    Two types of fixed pad hydrodynamic bearings (multilobe and pressure dam) were considered. Optimum and nonoptimum geometric configurations were tested. The optimum geometric configurations were determined by using a theoretical analysis and then the bearings were constructed for a flexible rotor test rig. It was found that optimizing bearings using this technique produces a 100% or greater increase in rotor stability. It is shown that this increase in rotor stability is carried out in the absence of certain types of instability mechanisms such as aerodynamic crosscoupling. However, the increase in rotor stability should greatly improve rotating machinery performance in the presence of such forces as well.

  1. Effects of a Rotating Aerodynamic Probe on the Flow Field of a Compressor Rotor

    NASA Technical Reports Server (NTRS)

    Lepicovsky, Jan

    2008-01-01

    An investigation of distortions of the rotor exit flow field caused by an aerodynamic probe mounted in the rotor is described in this paper. A rotor total pressure Kiel probe, mounted on the rotor hub and extending up to the mid-span radius of a rotor blade channel, generates a wake that forms additional flow blockage. Three types of high-response aerodynamic probes were used to investigate the distorted flow field behind the rotor. These probes were: a split-fiber thermo-anemometric probe to measure velocity and flow direction, a total pressure probe, and a disk probe for in-flow static pressure measurement. The signals acquired from these high-response probes were reduced using an ensemble averaging method based on a once per rotor revolution signal. The rotor ensemble averages were combined to construct contour plots for each rotor channel of the rotor tested. In order to quantify the rotor probe effects, the contour plots for each individual rotor blade passage were averaged into a single value. The distribution of these average values along the rotor circumference is a measure of changes in the rotor exit flow field due to the presence of a probe in the rotor. These distributions were generated for axial flow velocity and for static pressure.

  2. Some aspects of the aeroacoustics of high-speed jets

    NASA Technical Reports Server (NTRS)

    Lighthill, James

    1993-01-01

    Some of the background to contemporary jet aeroacoustics is addressed. Then scaling laws for noise generation by low-Mach-number airflows and by turbulence convected at 'not so low' Mach number is reviewed. These laws take into account the influence of Doppler effects associated with the convection of aeroacoustic sources. Next, a uniformly valid Doppler-effect approximation exhibits the transition, with increasing Mach number of convection, from compact-source radiation at low Mach numbers to a statistical assemblage of conical shock waves radiated by eddies convected at supersonic speed. In jets, for example, supersonic eddy convection is typically found for jet exit speeds exceeding twice the atmospheric speed of sound. The Lecture continues by describing a new dynamical theory of the nonlinear propagation of such statistically random assemblages of conical shock waves. It is shown, both by a general theoretical analysis and by an illustrative computational study, how their propagation is dominated by a characteristic 'bunching' process. That process associated with a tendency for shock waves that have already formed unions with other shock waves to acquire an increased proneness to form further unions - acts so as to enhance the high-frequency part of the spectrum of noise emission from jets at these high exit speeds.

  3. An Overview of Computational Aeroacoustic Modeling at NASA Langley

    NASA Technical Reports Server (NTRS)

    Lockard, David P.

    2001-01-01

    The use of computational techniques in the area of acoustics is known as computational aeroacoustics and has shown great promise in recent years. Although an ultimate goal is to use computational simulations as a virtual wind tunnel, the problem is so complex that blind applications of traditional algorithms are typically unable to produce acceptable results. The phenomena of interest are inherently unsteady and cover a wide range of frequencies and amplitudes. Nonetheless, with appropriate simplifications and special care to resolve specific phenomena, currently available methods can be used to solve important acoustic problems. These simulations can be used to complement experiments, and often give much more detailed information than can be obtained in a wind tunnel. The use of acoustic analogy methods to inexpensively determine far-field acoustics from near-field unsteadiness has greatly reduced the computational requirements. A few examples of current applications of computational aeroacoustics at NASA Langley are given. There remains a large class of problems that require more accurate and efficient methods. Research to develop more advanced methods that are able to handle the geometric complexity of realistic problems using block-structured and unstructured grids are highlighted.

  4. Numerical Study of Aeroacoustic Sound on Performance of Bladeless Fan

    NASA Astrophysics Data System (ADS)

    Jafari, Mohammad; Sojoudi, Atta; Hafezisefat, Parinaz

    2017-03-01

    Aeroacoustic performance of fans is essential due to their widespread application. Therefore, the original aim of this paper is to evaluate the generated noise owing to different geometric parameters. In current study, effect of five geometric parameters was investigated on well performance of a Bladeless fan. Airflow through this fan was analyzed simulating a Bladeless fan within a 2 m×2 m×4 m room. Analysis of the flow field inside the fan and evaluating its performance were obtained by solving conservations of mass and momentum equations for aerodynamic investigations and FW-H noise equations for aeroacoustic analysis. In order to design Bladeless fan Eppler 473 airfoil profile was used as the cross section of this fan. Five distinct parameters, namely height of cross section of the fan, outlet angle of the flow relative to the fan axis, thickness of airflow outlet slit, hydraulic diameter and aspect ratio for circular and quadratic cross sections were considered. Validating acoustic code results, we compared numerical solution of FW-H noise equations for NACA0012 with experimental results. FW-H model was selected to predict the noise generated by the Bladeless fan as the numerical results indicated a good agreement with experimental ones for NACA0012. To validate 3-D numerical results, the experimental results of a round jet showed good agreement with those simulation data. In order to indicate the effect of each mentioned parameter on the fan performance, SPL and OASPL diagrams were illustrated.

  5. Aeroacoustic and aerodynamic applications of the theory of nonequilibrium thermodynamics

    NASA Technical Reports Server (NTRS)

    Horne, W. Clifton; Smith, Charles A.; Karamcheti, Krishnamurty

    1991-01-01

    Recent developments in the field of nonequilibrium thermodynamics associated with viscous flows are examined and related to developments to the understanding of specific phenomena in aerodynamics and aeroacoustics. A key element of the nonequilibrium theory is the principle of minimum entropy production rate for steady dissipative processes near equilibrium, and variational calculus is used to apply this principle to several examples of viscous flow. A review of nonequilibrium thermodynamics and its role in fluid motion are presented. Several formulations are presented of the local entropy production rate and the local energy dissipation rate, two quantities that are of central importance to the theory. These expressions and the principle of minimum entropy production rate for steady viscous flows are used to identify parallel-wall channel flow and irrotational flow as having minimally dissipative velocity distributions. Features of irrotational, steady, viscous flow near an airfoil, such as the effect of trailing-edge radius on circulation, are also found to be compatible with the minimum principle. Finally, the minimum principle is used to interpret the stability of infinitesimal and finite amplitude disturbances in an initially laminar, parallel shear flow, with results that are consistent with experiment and linearized hydrodynamic stability theory. These results suggest that a thermodynamic approach may be useful in unifying the understanding of many diverse phenomena in aerodynamics and aeroacoustics.

  6. Wind Tunnel Testing of a 6%-Scale Large Civil Tilt Rotor Model in Airplane and Helicopter Modes

    DTIC Science & Technology

    2014-01-01

    were taken using the wind tunnel scale system. In addition to force and moment measurements, flow visualization using tufts, infrared thermography ... build -up and model reconfiguration time, and ultimately reduced the cost of the test entry. In particular, the HETR wing and nacelle geometries were...shared between the two models, which reduced fabrication, build -up and model reconfiguration time, and ultimately reduced the cost of the test entry

  7. Effect of AFT Rotor on the Inter-Rotor Flow of an Open Rotor Propulsion System

    NASA Technical Reports Server (NTRS)

    Slaboch, Paul E.; Stephens, David B.; Van Zante, Dale E.

    2016-01-01

    The effects of the aft rotor on the inter-rotor flow field of an open rotor propulsion rig were examined. A Particle Image Velocimetry (PIV) dataset that was acquired phase locked to the front rotor position has been phase averaged based on the relative phase angle between the forward and aft rotors. The aft rotor phase was determined by feature tracking in raw PIV images through an image processing algorithm. The effect of the aft rotor potential field on the inter-rotor flow were analyzed and shown to be in good agreement with Computational Fluid Dynamics (CFD) simulations. It was shown that the aft rotor had no substantial effect on the position of the forward rotor tip vortex but did have a small effect on the circulation strength of the vortex when the rotors were highly loaded.

  8. AMELIA Tests in NASA Wind Tunnel

    NASA Video Gallery

    This report from "This Week @ NASA" describes recent aerodynamic tests of a subscale model of the Advanced Model for Extreme Lift and Improved Aeroacoustics, or "AMELIA," in a NASA wind tunnel. The...

  9. Engine Rotor Dynamics, Synchronous and Nonsynchronous Whirl Control

    DTIC Science & Technology

    1979-02-01

    safely through or near the flexural critical speeds, a rotor shaft must either be well balanced , highly damped, or both. Nonsynchronous whirl induoed... rotor . The shaft is provided with five balance planes. This allows for multiplane balanc ’! of the rig as well as provides controlled imbalance. The...supercritical blade loss test in that the rotor was not rebalanced by jettisoning a second balance weight before deceleration. Figure 40 shows the midshaft

  10. Effect of higher harmonic control on helicopter rotor blade-vortex interaction noise: Prediction and initial validation

    NASA Technical Reports Server (NTRS)

    Beaumier, P.; Prieur, J.; Rahier, G.; Spiegel, P.; Demargne, A.; Tung, C.; Gallman, J. M.; Yu, Y. H.; Kube, R.; Vanderwall, B. G.

    1995-01-01

    The paper presents a status of theoretical tools of AFDD, DLR, NASA and ONERA for prediction of the effect of HHC on helicopter main rotor BVI noise. Aeroacoustic predictions from the four research centers, concerning a wind tunnel simulation of a typical descent flight case without and with HHC are presented and compared. The results include blade deformation, geometry of interacting vortices, sectional loads and noise. Acoustic predictions are compared to experimental data. An analysis of the results provides a first insight of the mechanisms by which HHC may affect BVI noise.

  11. ICASE/LaRC Workshop on Benchmark Problems in Computational Aeroacoustics (CAA)

    NASA Technical Reports Server (NTRS)

    Hardin, Jay C. (Editor); Ristorcelli, J. Ray (Editor); Tam, Christopher K. W. (Editor)

    1995-01-01

    The proceedings of the Benchmark Problems in Computational Aeroacoustics Workshop held at NASA Langley Research Center are the subject of this report. The purpose of the Workshop was to assess the utility of a number of numerical schemes in the context of the unusual requirements of aeroacoustical calculations. The schemes were assessed from the viewpoint of dispersion and dissipation -- issues important to long time integration and long distance propagation in aeroacoustics. Also investigated were the effect of implementation of different boundary conditions. The Workshop included a forum in which practical engineering problems related to computational aeroacoustics were discussed. This discussion took the form of a dialogue between an industrial panel and the workshop participants and was an effort to suggest the direction of evolution of this field in the context of current engineering needs.

  12. Testing and evaluation of a stall-flutter-suppression system for helicopter rotors using individual-blade-control

    NASA Technical Reports Server (NTRS)

    Quackenbush, T. R.

    1981-01-01

    The development and testing of a feedback system designed to alleviate the violent blade first torsion mode oscillations associated with stall flutter are described. The system, based on previously developed M.I.T. Individual-Blade-Control hardware, employs blade-mounted accelerometers to sense torsional oscillations and feeds back rate informaton to increase the damping of the first torsion mode. A linear model of the blade and control system dynamics is developed and is used to give qualitative and quantitative guidance in the design process as well as to aid in analysis of experimental results. System performance in wind tunnel tests, both in hover and forward flight, is described, and evidence is given of the system's ability to provide substantial additional damping to stall-induced blade oscillations.

  13. A Superior Kirchhoff Method for Aeroacoustic Noise Prediction: The Ffowcs Williams-Hawkings Equation

    NASA Technical Reports Server (NTRS)

    Brentner, Kenneth S.

    1997-01-01

    The prediction of aeroacoustic noise is important; all new aircraft must meet noise certification requirements. Local noise standards can be even more stringent. The NASA noise reduction goal is to reduce perceived noise levels by a factor of two in 10 years. The objective of this viewgraph presentation is to demonstrate the superiority of the FW-H approach over the Kirchoff method for aeroacoustics, both analytically and numerically.

  14. HPOTP low-speed flexible rotor balancing, phase 1

    NASA Technical Reports Server (NTRS)

    Giordano, J.; Zorzi, E.

    1985-01-01

    A method was developed that shows promise in overcoming many balancing limitations. This method establishes one or more windows for low speed, out-of-housing balancing of flexible rotors. These windows are regions of speed and support flexibility where two conditions are simultaneously fulfilled. First, the rotor system behaves flexibly; therefore, there is separation among balance planes. Second, the response due to balance weights is large enough to reliably measure. The analytic formulation of the low-speed flexible rotor balancing method is described. The results of proof-of-principle tests conducted under the program are presented. Based on this effort, it is concluded that low speed flexible rotor balancing is a viable technology. In particular, the method can be used to balance a rotor bearing system at low speed which results in smooth operation above more than one bending critical speed. Furthermore, this balancing methodology is applicable to SSME turbopump rotors.

  15. Aeroacoustic performance of a scoop inlet

    NASA Technical Reports Server (NTRS)

    Abbott, J. M.

    1977-01-01

    A low speed wind tunnel test demonstrated the aerodynamic and acoustic performance of a scoop inlet. Engine noise is directed upward by the extended lower lip of the scoop inlet. In addition, more of the scoop airflow comes in from above the inlet than below, leading to relatively higher surface velocities on the upper lip and lower surface velocities on the lower lip. These lower velocities on the lower lip result in a higher attainable angle of attack before internal flow separation occurs.

  16. A rotor technology assessment of the advancing blade concept

    NASA Technical Reports Server (NTRS)

    Pleasants, W. A.

    1983-01-01

    A rotor technology assessment of the Advancing Blade Concept (ABC) was conducted in support of a preliminary design study. The analytical methodology modifications and inputs, the correlation, and the results of the assessment are documented. The primary emphasis was on the high-speed forward flight performance of the rotor. The correlation data base included both the wind tunnel and the flight test results. An advanced ABC rotor design was examined; the suitability of the ABC for a particular mission was not considered. The objective of this technology assessment was to provide estimates of the performance potential of an advanced ABC rotor designed for high speed forward flight.

  17. Numerical simulation of a hovering rotor using embedded grids

    NASA Technical Reports Server (NTRS)

    Duque, Earl-Peter N.; Srinivasan, Ganapathi R.

    1992-01-01

    The flow field for a rotor blade in hover was computed by numerically solving the compressible thin-layer Navier-Stokes equations on embedded grids. In this work, three embedded grids were used to discretize the flow field - one for the rotor blade and two to convect the rotor wake. The computations were performed at two hovering test conditions, for a two-bladed rectangular rotor of aspect ratio six. The results compare fairly with experiment and illustrates the use of embedded grids in solving helicopter type flow fields.

  18. Aeroacoustics research in Europe: The CEAS-ASC report on 2014 highlights

    NASA Astrophysics Data System (ADS)

    Detandt, Yves

    2015-11-01

    The Council of European Aerospace Societies (CEAS) Aeroacoustics Specialists Committee (ASC) supports and promotes the interests of the scientific and industrial aeroacoustics community on an European scale and European aeronautics activities internationally. Each year the committee highlights some of the research and development projects in Europe. This paper is the 2014 issue of this collection of Aeroacoustic Highlights, compiled from informations submitted to the CEAS-ASC. The contributions are classified in different topics; the first categories being related to specific aeroacoustic challenges (airframe noise, fan and jet noise, helicopter noise, aircraft interior noise) and two last sections are respectively devoted to recent improvements and emerging techniques and to general advances in aeroacoustics. For each section, the present paper focus on accomplished projects, providing the state of the art in each research category in 2014. A number of research programmes involving aeroacoustics were funded by the European Commission. Some of the highlights from these programmes are summarised in this paper, as well as highlights funded by national programmes or by industry.

  19. Loads and performance data from a wind-tunnel test of model articulated helicopter rotors with 2 different blade torsional stiffnesses

    NASA Technical Reports Server (NTRS)

    Yeager, W. T., Jr.; Mantay, W. R.

    1983-01-01

    A passive means of tailoring helicopter rotor blades to improve performance and reduce loads was evaluated. The parameters investigated were blade torsional stiffness, blade section camber, and distance between blade structural elastic axis and blade tip aerodynamic center. This offset was accomplished by sweeping the tip. The investigation was conducted at advance ratios of 0.20, 0.30, and 0.40. Data are presented without analysis; however, cross referencing of performance data and harmonic loads data may be useful to the analyst for validating aeroelastic theories and design methodologies as well as for evaluating passive aeroelastic tailoring or rotor blade parameters.

  20. Rotor blade vortex interaction noise

    NASA Astrophysics Data System (ADS)

    Yu, Yung H.

    2000-02-01

    Blade-vortex interaction noise-generated by helicopter main rotor blades is one of the most severe noise problems and is very important both in military applications and community acceptance of rotorcraft. Research over the decades has substantially improved physical understanding of noise-generating mechanisms, and various design concepts have been investigated to control noise radiation using advanced blade planform shapes and active blade control techniques. The important parameters to control rotor blade-vortex interaction noise and vibration have been identified: blade tip vortex structures and its trajectory, blade aeroelastic deformation, and airloads. Several blade tip design concepts have been investigated for diffusing tip vortices and also for reducing noise. However, these tip shapes have not been able to substantially reduce blade-vortex interaction noise without degradation of rotor performance. Meanwhile, blade root control techniques, such as higher-harmonic pitch control (HHC) and individual blade control (IBC) concepts, have been extensively investigated for noise and vibration reduction. The HHC technique has proved the substantial blade-vortex interaction noise reduction, up to 6 dB, while vibration and low-frequency noise have been increased. Tests with IBC techniques have shown the simultaneous reduction of rotor noise and vibratory loads with 2/rev pitch control inputs. Recently, active blade control concepts with smart structures have been investigated with the emphasis on active blade twist and trailing edge flap. Smart structures technologies are very promising, but further advancements are needed to meet all the requirements of rotorcraft applications in frequency, force, and displacement.

  1. An outflow boundary condition for aeroacoustic computations

    NASA Technical Reports Server (NTRS)

    Hayder, M. Ehtesham; Hagstrom, Thomas

    1995-01-01

    A formulation of boundary condition for flows with small disturbances is presented. The authors test their methodology in an axisymmetric jet flow calculation, using both the Navier-Stokes and Euler equations. Solutions in the far field are assumed to be oscillatory. If the oscillatory disturbances are small, the growth of the solution variables can be predicted by linear theory. Eigenfunctions of the linear theory are used explicitly in the formulation of the boundary conditions. This guarantees correct solutions at the boundary in the limit where the predictions of linear theory are valid.

  2. Aeroacoustic performance of a scoop inlet

    NASA Technical Reports Server (NTRS)

    Abbott, J. M.

    1977-01-01

    Results of a low speed wind tunnel test program are presented which demonstrate the aerodynamic and acoustic performance of a scoop inlet. Engine noise that would normally propagate toward the ground is directed upward by the extended lower lip of the scoop inlet. In addition, more of the scoop airflow comes in from above the inlet than below, leading to relatively higher surface velocities on the upper lip and lower surface velocities on the lower lip. These lower velocities on the lower lip result in a higher attainable angle of attack before internal flow separation occurs.

  3. Aeroacoustic wind tunnel measurements on propeller noise

    NASA Astrophysics Data System (ADS)

    Grosche, F. R.; Stiewitt, H.

    1985-02-01

    Model tests were conducted in a low speed wind tunnel to determine the sound radiation of 5 propellers with different blade designs including variations of thickness ratios, blade profiles, blade planforms and blade tip configurations. The diameter of the propellers was 0.9 m, the propeller speed was kept constant. The tip Mach number was M sub I = 0.66 and the helical tip Mach number varied between 0.66 and 0.69. The main objectives were to investigate the effects of blade geometry on near field and far field noise and to locate the dominant sound sources in the propeller plane, radiating to the observer, by means of a highly directional microphone system. The results include: (1) comparisons of noise spectra of different propeller configurations; (2) near field sound pressures as function of axial distance from the propeller plane; and (3) directivity of sound radiation from the moving blades.

  4. Development of Experimental and Computational Aeroacoustic Tools for Advanced Liner Evaluation

    NASA Technical Reports Server (NTRS)

    Jones, Michael G.; Watson, Willie R.; Nark, Douglas N.; Parrott, Tony L.; Gerhold, Carl H.; Brown, Martha C.

    2006-01-01

    Acoustic liners in aircraft engine nacelles suppress radiated noise. Therefore, as air travel increases, increasingly sophisticated tools are needed to maximize noise suppression. During the last 30 years, NASA has invested significant effort in development of experimental and computational acoustic liner evaluation tools. The Curved Duct Test Rig is a 152-mm by 381- mm curved duct that supports liner evaluation at Mach numbers up to 0.3 and source SPLs up to 140 dB, in the presence of user-selected modes. The Grazing Flow Impedance Tube is a 51- mm by 63-mm duct currently being fabricated to operate at Mach numbers up to 0.6 with source SPLs up to at least 140 dB, and will replace the existing 51-mm by 51-mm duct. Together, these test rigs allow evaluation of advanced acoustic liners over a range of conditions representative of those observed in aircraft engine nacelles. Data acquired with these test ducts are processed using three aeroacoustic propagation codes. Two are based on finite element solutions to convected Helmholtz and linearized Euler equations. The third is based on a parabolic approximation to the convected Helmholtz equation. The current status of these computational tools and their associated usage with the Langley test rigs is provided.

  5. Redesigned rotor for a highly loaded, 1800 ft/sec tip speed compressor fan stage 1: Aerodynamic and mechanical design

    NASA Technical Reports Server (NTRS)

    Halle, J. E.; Ruschak, J. T.

    1975-01-01

    A highly loaded, high tip-speed fan rotor was designed with multiple-circular-arc airfoil sections as a replacement for a marginally successful rotor which had precompression airfoil sections. The substitution of airfoil sections was the only aerodynamic change. Structural design of the redesigned rotor blade was guided by successful experience with the original blade. Calculated stress levels and stability parameters for the redesigned rotor are within limits demonstrated in tests of the original rotor.

  6. Development of a micromachined piezoelectric microphone for aeroacoustics applications.

    PubMed

    Horowitz, Stephen; Nishida, Toshikazu; Cattafesta, Louis; Sheplak, Mark

    2007-12-01

    This paper describes the design, fabrication, and characterization of a bulk-micromachined piezoelectric microphone for aeroacoustic applications. Microphone design was accomplished through a combination of piezoelectric composite plate theory and lumped element modeling. The device consists of a 1.80-mm-diam, 3-microm-thick, silicon diaphragm with a 267-nm-thick ring of piezoelectric material placed near the boundary of the diaphragm to maximize sensitivity. The microphone was fabricated by combining a sol-gel lead zirconate-titanate deposition process on a silicon-on-insulator wafer with deep-reactive ion etching for the diaphragm release. Experimental characterization indicates a sensitivity of 1.66 microVPa, dynamic range greater than six orders of magnitude (35.7-169 dB, re 20 microPa), a capacitance of 10.8 nF, and a resonant frequency of 59.0 kHz.

  7. An Aeroacoustic Study of a Leading Edge Slat Configuration

    NASA Technical Reports Server (NTRS)

    Mendoza, J. M.; Brooks, T. F.; Humphreys, W. M., Jr.

    2002-01-01

    Aeroacoustic evaluations of high-lift devices have been carried out in the Quiet Flow Facility of the NASA Langley Research Center. The present paper describes detailed flow and acoustic measurements that have been made in order to better understand the noise generated from airflow over a wing leading edge slat configuration, and to possibly predict and reduce this noise source. The acoustic database is obtained by a moveable Small Aperture Directional Array of microphones designed to electronically steer to different portions of models under study. The slat is shown to be a uniform distributed noise source. The data was processed such that spectra and directivity were determined with respect to a one-foot span of slat. The spectra are normalized in various fashions to demonstrate slat noise character. In order to equate portions of the spectra to different slat noise components, trailing edge noise predictions using measured slat boundary layer parameters as inputs are compared to the measured slat noise spectra.

  8. A general introduction to aeroacoustics and atmospheric sound

    NASA Technical Reports Server (NTRS)

    Lighthill, James

    1992-01-01

    A single unifying principle (based upon the nonlinear 'momentum-flux' effects produced when different components of a motion transport different components of its momentum) is used to give a broad scientific background to several aspects of the interaction between airflows and atmospheric sound. First, it treats the generation of sound by airflows of many different types. These include, for example, jet-like flows involving convected turbulent motions (with the resulting aeroacoustic radiation sensitively dependent on the Mach number of convection) and they include, as an extreme case, the supersonic 'boom' (shock waves generated by a supersonically convected flow pattern). Next, an analysis is given of sound propagation through nonuniformly moving airflows, and the exchange is quantified of energy between flow and sound; while, finally, problems are examined of how sound waves 'on their own' may generate the airflows known as acoustic streaming.

  9. Computational aeroacoustics applications based on a discontinuous Galerkin method

    NASA Astrophysics Data System (ADS)

    Delorme, Philippe; Mazet, Pierre; Peyret, Christophe; Ventribout, Yoan

    2005-09-01

    CAA simulation requires the calculation of the propagation of acoustic waves with low numerical dissipation and dispersion error, and to take into account complex geometries. To give, at the same time, an answer to both challenges, a Discontinuous Galerkin Method is developed for Computational AeroAcoustics. Euler's linearized equations are solved with the Discontinuous Galerkin Method using flux splitting technics. Boundary conditions are established for rigid wall, non-reflective boundary and imposed values. A first validation, for induct propagation is realized. Then, applications illustrate: the Chu and Kovasznay's decomposition of perturbation inside uniform flow in term of independent acoustic and rotational modes, Kelvin-Helmholtz instability and acoustic diffraction by an air wing. To cite this article: Ph. Delorme et al., C. R. Mecanique 333 (2005).

  10. Navier-Stokes computations of cavity aeroacoustics with suppression devices

    NASA Technical Reports Server (NTRS)

    Baysal, Oktay; Yen, Guan-Wei; Fouladi, Kamran

    1992-01-01

    Effectiveness of two devices to suppress the cavity acoustics was computationally investigated. Two dimensional, computational simulations were performed for the transonic, turbulent flows past a cavity, which was first equipped with a rear face ramp and then with a spoiler. The Reynolds-averaged, unsteady, compressible, full Navier-Stokes equations were solved time accurately by a second-order accurate, implicit, upwind, finite-volume method. The effect of turbulence was included through a Baldwin-Lomax model with modifications for the multiple-wall effects and for the highly vortical flow with a shear layer. The results included instantaneous and time-averaged flow properties, and time-series analyses of the pressure inside the cavity, which compared favorably with the available experimental data. These results were also contrasted with the computed aeroacoustics of the same cavity (length-to-depth ratio of 4.5), but without a device, to demonstrate the suppression effectiveness.

  11. Benchmark Solutions for Computational Aeroacoustics (CAA) Code Validation

    NASA Technical Reports Server (NTRS)

    Scott, James R.

    2004-01-01

    NASA has conducted a series of Computational Aeroacoustics (CAA) Workshops on Benchmark Problems to develop a set of realistic CAA problems that can be used for code validation. In the Third (1999) and Fourth (2003) Workshops, the single airfoil gust response problem, with real geometry effects, was included as one of the benchmark problems. Respondents were asked to calculate the airfoil RMS pressure and far-field acoustic intensity for different airfoil geometries and a wide range of gust frequencies. This paper presents the validated that have been obtained to the benchmark problem, and in addition, compares them with classical flat plate results. It is seen that airfoil geometry has a strong effect on the airfoil unsteady pressure, and a significant effect on the far-field acoustic intensity. Those parts of the benchmark problem that have not yet been adequately solved are identified and presented as a challenge to the CAA research community.

  12. High Order Difference Method for Low Mach Number Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Mueller, B.; Yee, H. C.; Mansour, Nagi (Technical Monitor)

    2001-01-01

    A high order finite difference method with improved accuracy and stability properties for computational aeroacoustics (CAA) at low Mach numbers is proposed. The Euler equations are split into a conservative and a symmetric non- conservative portion to allow the derivation of a generalized energy estimate. Since the symmetrization is based on entropy variables, that splitting of the flux derivatives is referred to as entropy splitting. Its discretization by high order central differences was found to need less numerical dissipation than conventional conservative schemes. Owing to the large disparity of acoustic and stagnation quantities in low Mach number aeroacoustics, the split Euler equations are formulated in perturbation form. The unknowns are the small changes of the conservative variables with respect to their large stagnation values. All nonlinearities and the conservation form of the conservative portion of the split flux derivatives can be retained, while cancellation errors are avoided with its discretization opposed to the conventional conservative form. The finite difference method is third-order accurate at the boundary and the conventional central sixth-order accurate stencil in the interior. The difference operator satisfies the summation by parts property analogous to the integration by parts in the continuous energy estimate. Thus, strict stability of the difference method follows automatically. Spurious high frequency oscillations are suppressed by a characteristic-based filter similar to but without limiter. The time derivative is approximated by a 4-stage low-storage second-order explicit Runge-Kutta method. The method has been applied to simulate vortex sound at low Mach numbers. We consider the Kirchhoff vortex, which is an elliptical patch of constant vorticity rotating with constant angular frequency in irrotational flow. The acoustic pressure generated by the Kirchhoff vortex is governed by the 2D Helmholtz equation, which can be solved

  13. Aeroacoustic prediction of turbulent free shear flows

    NASA Astrophysics Data System (ADS)

    Bodony, Daniel Joseph

    2005-12-01

    For many people living in the immediate vicinity of an active airport the noise of jet aircraft flying overhead can be a nuisance, if not worse. Airports, which are held accountable for the noise they produce, and upcoming international noise limits are pressuring the major airframe and jet engine manufacturers to bring quieter aircraft into service. However, component designers need a predictive tool that can estimate the sound generated by a new configuration. Current noise prediction techniques are almost entirely based on previously collected experimental data and are applicable only to evolutionary, not revolutionary, changes in the basic design. Physical models of final candidate designs must still be built and tested before a single design is selected. By focusing on the noise produced in the jet engine exhaust at take-off conditions, the prediction of sound generated by turbulent flows is addressed. The technique of large-eddy simulation is used to calculate directly the radiated sound produced by jets at different operating conditions. Predicted noise spectra agree with measurements for frequencies up to, and slightly beyond, the peak frequency. Higher frequencies are missed, however, due to the limited resolution of the simulations. Two methods of estimating the 'missing' noise are discussed. In the first a subgrid scale noise model, analogous to a subgrid scale closure model, is proposed. In the second method the governing equations are expressed in a wavelet basis from which simplified time-dependent equations for the subgrid scale fluctuations can be derived. These equations are inexpensively integrated to yield estimates of the subgrid scale fluctuations with proper space-time dynamics.

  14. An investigation of rotor harmonic noise by the use of small scale wind tunnel models

    NASA Technical Reports Server (NTRS)

    Sternfeld, H., Jr.; Schaffer, E. G.

    1982-01-01

    Noise measurements of small scale helicopter rotor models were compared with noise measurements of full scale helicopters to determine what information about the full scale helicopters could be derived from noise measurements of small scale helicopter models. Comparisons were made of the discrete frequency (rotational) noise for 4 pairs of tests. Areas covered were tip speed effects, isolated rotor, tandem rotor, and main rotor/tail rotor interaction. Results show good comparison of noise trends with configuration and test condition changes, and good comparison of absolute noise measurements with the corrections used except for the isolated rotor case. Noise measurements of the isolated rotor show a great deal of scatter reflecting the fact that the rotor in hover is basically unstable.

  15. Boost pulverizer performance with new exhauster rotor

    SciTech Connect

    Lauber, J.A.

    1994-10-01

    Georgia Power Co.'s Jack McDonough station, near Atlanta, Ga, consists of two tangentially fired boilers rated 1.734-million lb/hr, each paired with a 265-MW turbine/generator. The plant's primary fuel is Eastern Bituminous coal, which is ground by five pulverizers. In April 1993, it first began evaluating a new high-efficiency rotor for the pulverizer exhausters, supplied by ABB C-E Services Inc. Windsor, Conn. McDonough was designed to burn coal with a Hardgrove Grindability Index (HGI) of 55 and is typically supplied fuel with an HGI of 45. Thus, the units must operate with all five pulverizers to obtain full load, although each was intended to generate full load using four pulverizers. Now that the plant is burning low-sulfur coal to comply with the clean Air Act Amendments of 1990, the increased hardness limits McDonough's full-load capabilities and will dramatically increase pulverizer maintenance. Management was left with two options: a complete pulverizer replacement or further modifications. (The mills had already been upgraded several years ago.) The introduction of high-efficiency rotors for the exhauster fans made the second option more attractive. Unfortunately, the vendor had no significant operating experience with the rotor design, still in the developmental stage. Therefore, the plant decided to install a test rotor at McDonough Unit 2. This article is a review of the rotor's basic theory of operation, test results, and wear characteristics experienced at the unit.

  16. Field Balancing and Harmonic Vibration Suppression in Rigid AMB-Rotor Systems with Rotor Imbalances and Sensor Runout.

    PubMed

    Xu, Xiangbo; Chen, Shao

    2015-08-31

    Harmonic vibrations of high-speed rotors in momentum exchange devices are primary disturbances for attitude control of spacecraft. Active magnetic bearings (AMBs), offering the ability to control the AMB-rotor dynamic behaviors, are preferred in high-precision and micro-vibration applications, such as high-solution Earth observation satellites. However, undesirable harmonic displacements, currents, and vibrations also occur in the AMB-rotor system owing to the mixed rotor imbalances and sensor runout. To compensate the rotor imbalances and to suppress the harmonic vibrations, two control methods are presented. Firstly, a four degrees-of-freedom AMB-rotor model with the static imbalance, dynamic imbalance, and the sensor runout are described. Next, a synchronous current reduction approach with a variable-phase notch feedback is proposed, so that the rotor imbalances can be identified on-line through the analysis of the synchronous displacement relationships of the geometric, inertial, and rotational axes of the rotor. Then, the identified rotor imbalances, which can be represented at two prescribed balancing planes of the rotor, are compensated by discrete add-on weights whose masses are calculated in the vector form. Finally, a repetitive control algorithm is utilized to suppress the residual harmonic vibrations. The proposed field balancing and harmonic vibration suppression strategies are verified by simulations and experiments performed on a control moment gyro test rig with a rigid AMB-rotor system. Compared with existing methods, the proposed strategies do not require trial weights or an accurate model of the AMB-rotor system. Moreover, the harmonic displacements, currents, and vibrations can be well-attenuated simultaneously.

  17. Field Balancing and Harmonic Vibration Suppression in Rigid AMB-Rotor Systems with Rotor Imbalances and Sensor Runout

    PubMed Central

    Xu, Xiangbo; Chen, Shao

    2015-01-01

    Harmonic vibrations of high-speed rotors in momentum exchange devices are primary disturbances for attitude control of spacecraft. Active magnetic bearings (AMBs), offering the ability to control the AMB-rotor dynamic behaviors, are preferred in high-precision and micro-vibration applications, such as high-solution Earth observation satellites. However, undesirable harmonic displacements, currents, and vibrations also occur in the AMB-rotor system owing to the mixed rotor imbalances and sensor runout. To compensate the rotor imbalances and to suppress the harmonic vibrations, two control methods are presented. Firstly, a four degrees-of-freedom AMB-rotor model with the static imbalance, dynamic imbalance, and the sensor runout are described. Next, a synchronous current reduction approach with a variable-phase notch feedback is proposed, so that the rotor imbalances can be identified on-line through the analysis of the synchronous displacement relationships of the geometric, inertial, and rotational axes of the rotor. Then, the identified rotor imbalances, which can be represented at two prescribed balancing planes of the rotor, are compensated by discrete add-on weights whose masses are calculated in the vector form. Finally, a repetitive control algorithm is utilized to suppress the residual harmonic vibrations. The proposed field balancing and harmonic vibration suppression strategies are verified by simulations and experiments performed on a control moment gyro test rig with a rigid AMB-rotor system. Compared with existing methods, the proposed strategies do not require trial weights or an accurate model of the AMB-rotor system. Moreover, the harmonic displacements, currents, and vibrations can be well-attenuated simultaneously. PMID:26334281

  18. SMART wind turbine rotor. Data analysis and conclusions

    SciTech Connect

    Berg, Jonathan Charles; Barone, Matthew Franklin; Yoder, Nathanael C.

    2014-01-01

    The Wind Energy Technologies department at Sandia National Laboratories has developed and field tested a wind turbine rotor with integrated trailing-edge flaps designed for active control of the rotor aerodynamics. The SMART Rotor project was funded by the Wind and Water Power Technologies Office of the U.S. Department of Energy (DOE) and was conducted to demonstrate active rotor control and evaluate simulation tools available for active control research. This report documents the data post-processing and analysis performed to date on the field test data. Results include the control capability of the trailing edge flaps, the combined structural and aerodynamic damping observed through application of step actuation with ensemble averaging, direct observation of time delays associated with aerodynamic response, and techniques for characterizing an operating turbine with active rotor control.

  19. Acoustic Aspects of Active-Twist Rotor Control

    NASA Technical Reports Server (NTRS)

    Booth, Earl R., Jr.; Wilbur, Matthew L.

    2002-01-01

    The use of an Active Twist Rotor system to provide both vibration reduction and performance enhancement has been explored in recent analytical and experimental studies. Effects of active-twist control on rotor noise, however, had not been determined. During a recent wind tunnel test of an active-twist rotor system, a set of acoustic measurements were obtained to assess the effects of active-twist control on noise produced by the rotor, especially blade-vortex interaction (BVI) noise. It was found that for rotor operating conditions where BVI noise is dominant, active-twist control provided a reduction in BVI noise level. This BVI noise reduction was almost, but not quite, as large as that obtained in a similar test using HHC. However, vibration levels were usually adversely affected at operating conditions favoring minimum BVI noise. Conversely, operating conditions favoring minimum vibration levels affected BVI noise levels, but not always adversely.

  20. Towards an effective non-reflective boundary condition for computational aeroacoustics

    NASA Astrophysics Data System (ADS)

    Gill, James; Fattah, Ryu; Zhang, Xin

    2017-03-01

    A generic, non-reflective zonal transverse characteristic boundary condition is described for computational aeroacoustics, which shows superior performance to existing non-reflective boundary conditions for two-dimensional linearized Euler simulations. The new condition is based on a characteristic non-reflective method, and also contains optimised use of transverse characteristic terms and a zonal forcing region. The performance of the new method and several existing non-reflective acoustic boundary conditions is quantitatively compared using a plane wave test case. The performance of buffer zone, perfectly matched layer, far-field, and characteristic non-reflective methods is compared, following an optimisation of the tuneable parameters in each method to give best performance. The study uses a high-order linearised Euler equation solver to assess non-reflective boundary conditions with a variety of cases. The performance is compared for downstream travelling acoustic waves with varying frequency and incident angle, and at various Mach numbers. The current study includes a more comprehensive evaluation than previous studies which used constant values of tuneable parameters or qualitative assessment methods. The new zonal transverse characteristic boundary condition is shown to give improved performance in comparison to the other tested outflow boundary conditions for two-dimensional linearized Euler simulations, and is also shown to give good performance when used as an inflow condition.

  1. Rotor balancing apparatus and system

    NASA Technical Reports Server (NTRS)

    Lyman, Frank (Inventor); Lyman, Joseph (Inventor)

    1976-01-01

    Rotor balancing apparatus and a system comprising balance probes for measuring unbalance at the ends of a magnetically suspended rotor are disclosed. Each balance probe comprises a photocell which is located in relationship to the magnetically suspended rotor such that unbalance of the rotor changes the amount of light recorded by each photocell. The signal from each photocell is electrically amplified and displayed by a suitable device, such as an oscilloscope.

  2. Helicopter tail rotor noise analyses

    NASA Technical Reports Server (NTRS)

    George, A. R.; Chou, S. T.

    1986-01-01

    A study was made of helicopter tail rotor noise, particularly that due to interactions with the main rotor tip vortices, and with the fuselage separation mean wake. The tail rotor blade-main rotor tip vortex interaction is modelled as an airfoil of infinite span cutting through a moving vortex. The vortex and the geometry information required by the analyses are obtained through a free wake geometry analysis of the main rotor. The acoustic pressure-time histories for the tail rotor blade-vortex interactions are then calculated. These acoustic results are compared to tail rotor loading and thickness noise, and are found to be significant to the overall tail rotor noise generation. Under most helicopter operating conditions, large acoustic pressure fluctuations can be generated due to a series of skewed main rotor tip vortices passing through the tail rotor disk. The noise generation depends strongly upon the helicopter operating conditions and the location of the tail rotor relative to the main rotor.

  3. Rotor/bearing system dynamic stiffness measurements

    NASA Technical Reports Server (NTRS)

    Muszynska, A.

    1985-01-01

    Sweep perturbation testing as used in Modal Analysis when applied to a rotating machine has to take into consideration the machine dynamic state of equilibrium at its operational rotative speed. This stands in contrasts to a static equilibrium of nonrotating structures. The rotational energy has a significant influence on rotor dynamic characteristics. The best perturbing input for rotating machines is a forward or reverse rotating, circular force applied directly to the shaft. Determination of Dynamic Stiffness Characteristics of the rotor bearing system by nonsynchronous perturbation of a symmetric rotating shaft supported in one relatively rigid and one oil lubricated bearing.

  4. Higher harmonic rotor blade pitch control

    NASA Technical Reports Server (NTRS)

    Ewans, J. R.

    1976-01-01

    Tests of a model 'Reverse Velocity Rotor' system at high advance ratios and with twice-per-revolution cyclic pitch control were made under joint Navy-NASA sponsorship in the NASA, Ames 12 ft. pressure tunnel. The results showed significant gains in rotor performance at all advance ratios by using twice-per-revolution control. Detailed design studies have been made of alternative methods of providing higher harmonic motion including four types of mechanical systems and an electro-hydraulic system. The relative advantages and disadvantages are evaluated on the basis of stiffness, weight, volume, reliability and maintainability.

  5. Application of High-Temperature Mold Materials to Die Cast Copper Motor Rotor for Improved Efficiency

    SciTech Connect

    John G. Cowie; Edwin F. Brush, Jr.; Dale T. Peters; Stephen P. Midson; Darryl J. Van Son

    2003-05-01

    The objective of the study, Application of High-Temperature Mold Materials to Die Cast Copper Motor Rotor for Improved Efficiency, was to support the Copper Development Association (CDA) in its effort to design, fabricate and demonstrate mold technologies designed to withstand the copper motor rotor die casting environment for an economically acceptable life. The anticipated result from the compiled data and tests were to: (1) identify materials suitable for die casting copper, (2) fabricate motor rotor molds and (3) supply copper rotor motors for testing in actual compressor systems. Compressor manufacturers can apply the results to assess the technical and economical viability of copper rotor motors.

  6. Aeroacoustics research in Europe: The CEAS-ASC report on 2011 highlights

    NASA Astrophysics Data System (ADS)

    McAlpine, A.; Astley, R. J.

    2012-10-01

    The Council of European Aerospace Societies (CEAS) Aeroacoustics Specialists Committee (ASC) supports and promotes the interests of the scientific and industrial aeroacoustics community on an European scale and European aeronautics activities internationally. In this context, "aeroacoustics" encompasses all aerospace acoustics and related areas. Each year the committee highlights some of the research and development projects in Europe. This paper is a report on highlights of aeroacoustics research in Europe in 2011, compiled from information provided to the ASC of the CEAS. During 2011, a number of research programmes involving aeroacoustics were funded by the European Commission. Some of the highlights from these programmes are summarized in this paper, as well as highlights from other programmes funded by national programmes or by industry. Furthermore, a concise summary of the CEAS-ASC workshop "Acoustic Liners and Associated Propagation Techniques" held in Lausanne in October 2011 is included in this report. Enquiries concerning all contributions should be addressed to the authors who are given at the end of each subsection.

  7. Effect of rotor wake on aerodynamic characteristics of a 1/6 scale model of the rotor systems research aircraft. [in the Langley V/STOL tunnel

    NASA Technical Reports Server (NTRS)

    Mineck, R. E.

    1977-01-01

    Tests were conducted in the Langley V/STOL tunnel to determine the effect of the main-rotor wake on the aerodynamic characteristics of the rotor systems research aircraft. A 1/6-scale model with a 4-blade articulated rotor was used to determine the effect of the rotor wake for the compound configuration. Data were obtained over a range of angles of attack, angles of sideslip, auxiliary engine thrusts, rotor collective pitch angles, and rotor tip-path plane angles for several main-rotor advance ratios. Separate results are presented for the forces and moments on the airframe, the wing, and the tail. An analysis of the test data indicates significant changes in the aerodynamic characteristics. The rotor wake increases the longitudinal static stability, the effective dihedral, and the lateral static stability of the airframe. The rotor induces a downwash on the wing. This downwash decreases the wing lift and increases the drag. The asymmetrical rotor wake induces a differential lift across the wing and a subsequent rolling moment. These rotor induced effects on the wing become smaller with increasing forward speed.

  8. Performance of a tandem-rotor/tandem-stator conical-flow compressor designed for a pressure ratio of 3

    NASA Technical Reports Server (NTRS)

    Wood, J. R.; Owen, A. K.; Schumann, L. F.

    1982-01-01

    A conical-flow compressor stage with a large radius change through the rotor was tested at three values of rotor tip clearance. The stage had a tandem rotor and a tandem stator. Peak efficiency at design speed was 0.774 at a pressure ratio of 2.613. The rotor was tested without the stator, and detailed survey data were obtained for each rotor blade row. Overall peak rotor efficiency was 0.871 at a pressure ratio of 2.952.

  9. Wavenumber-frequency deconvolution of aeroacoustic microphone phased array data of arbitrary coherence

    NASA Astrophysics Data System (ADS)

    Bahr, Christopher J.; Cattafesta, Louis N.

    2016-11-01

    Deconvolution of aeroacoustic data acquired with microphone phased arrays is a computationally challenging task for distributed sources with arbitrary coherence. A new technique for performing such deconvolution is proposed. This technique relies on analysis of the array data in the wavenumber-frequency domain, allowing for fast convolution and reduced storage requirements when compared to traditional coherent deconvolution. A positive semidefinite constraint for the iterative deconvolution procedure is implemented and shows improved behavior in terms of quantifiable convergence metrics when compared to a standalone covariance inequality constraint. A series of simulations validates the method's ability to resolve coherence and phase angle relationships between partially coherent sources, as well as determines convergence criteria for deconvolution analysis. Simulations for point sources near the microphone phased array show potential for handling such data in the wavenumber-frequency domain. In particular, a physics-based integration boundary calculation is described, and can successfully isolate sources and track the appropriate integration bounds with and without the presence of flow. Magnitude and phase relationships between multiple sources are successfully extracted. Limitations of the deconvolution technique are determined from the simulations, particularly in the context of a simulated acoustic field in a closed test section wind tunnel with strong boundary layer contamination. A final application to a trailing edge noise experiment conducted in an open-jet wind tunnel matches best estimates of acoustic levels from traditional calculation methods and qualitatively assesses the coherence characteristics of the trailing edge noise source.

  10. Vibration and Control of Flexible Rotor Supported by Magnetic Bearings

    NASA Technical Reports Server (NTRS)

    Nonami, Kenzou

    1988-01-01

    Active vibration control of flexible rotors supported by magnetic bearings is discussed. Using a finite-element method for a mathematical model of the flexible rotor, the eigenvalue problem is formulated taking into account the interaction between a mechanical system of the flexible rotor and an electrical system of the magnetic bearings and the controller. However, for the sake of simplicity, gyroscopic effects are disregarded. It is possible to adapt this formulation to a general flexible rotor-magnetic bearing system. Controllability with and without collocation sensors and actuators located at the same distance along the rotor axis is discussed for the higher order flexible modes of the test rig. In conclusion, it is proposed that it is necessary to add new active control loops for the higher flexible modes even in the case of collocation. Then it is possible to stabilize for the case of uncollocation by means of this method.

  11. Analysis of a Hovering Rotor in Icing Conditions

    NASA Technical Reports Server (NTRS)

    Narducci, Robert; Kreeger, Richard E.

    2012-01-01

    A high fidelity analysis method is proposed to evaluate the ice accumulation and the ensuing rotor performance degradation for a helicopter flying through an icing cloud. The process uses computational fluid dynamics (CFD) coupled to a rotorcraft comprehensive code to establish the aerodynamic environment of a trimmed rotor prior to icing. Based on local aerodynamic conditions along the rotor span and accounting for the azimuthal variation, an ice accumulation analysis using NASA's Lewice3D code is made to establish the ice geometry. Degraded rotor performance is quantified by repeating the high fidelity rotor analysis with updates which account for ice shape and mass. The process is applied on a full-scale UH-1H helicopter in hover using data recorded during the Helicopter Icing Flight Test Program.

  12. Aerodynamic Evaluation of Two Compact Radial-Inflow Turbine Rotors

    NASA Technical Reports Server (NTRS)

    Simonyi, P. Susan; Roelke, Richard J.; Stabe, Roy G.; Nowlin, Brentley C.; Dicicco, Danielle

    1995-01-01

    The aerodynamic evaluation of two highly loaded compact radial turbine rotors was conducted at the NASA Lewis Research Center Small Engine Component Test Facility (SECTF). The experimental results were used for proof-of-concept, for modeling radial inflow turbine rotors, and for providing data for code verification. Two rotors were designed to have a shorter axial length, up to a 10-percent reduced diameter, a lighter weight, and equal or higher efficiencies with those of conventional radial inflow turbine rotors. Three configurations were tested: rotor 1, having a 40-percent shorter axial length, with the design stator (stator 1); rotor 1 with the design stator vanes closed down (stator 2); and rotor 2, slightly shorter axially and having higher loading, with stator 2. The stator had 36 vanes and the rotors each had 14 solid blades. Although presently uncooled, the rotor blades were designed for thicknesses which would allow cooling passages to be added. The overall stage performance measurements and the rotor and stator exit flow field surveys were obtained. Measurements of steady state temperatures, pressures, mass flow rates, flow angles, and output power were made at various operating conditions. Data were obtained at corrected speeds of 80, 90, 100, 110, and 120 percent of design over a range of equivalent inlet-to-exit pressure ratios of 3.5, 4.0, 4.5, and 5.0, the maximum pressure ratio achieved. The test showed that the configuration of rotor 1 with stator 1 running at the design pressure ratio produced a flow rate which was 5.6 percent higher than expected. This result indicated the need to close down the stator flow area to reduce the flow. The flow area reduction was accomplished by restaggering the vanes. Rotor 1 was retested with the closed-down stator vanes and achieved the correct mass flow. Rotor 2 was tested only with the restaggered vanes. The test results of the three turbine configurations were nearly identical. Although the measured

  13. Blended Wing Body Concept Development with Open Rotor Engine Intergration

    NASA Technical Reports Server (NTRS)

    Pitera, David M.; DeHaan, Mark; Brown, Derrell; Kawai, Ronald T.; Hollowell, Steve; Camacho, Peter; Bruns, David; Rawden, Blaine K.

    2011-01-01

    The purpose of this study is to perform a systems analysis of a Blended Wing Body (BWB) open rotor concept at the conceptual design level. This concept will be utilized to estimate overall noise and fuel burn performance, leveraging recent test data. This study will also investigate the challenge of propulsion airframe installation of an open rotor engine on a BWB configuration. Open rotor engines have unique problems relative to turbofans. The rotors are open, exposed to flow conditions outside of the engine. The flow field that the rotors are immersed in may be higher than the free stream flow and it may not be uniform, both of these characteristics could increase noise and decrease performance. The rotors sometimes cause changes in the flow conditions imposed on aircraft surfaces. At high power conditions such as takeoff and climb out, the stream tube of air that goes through the rotors contracts rapidly causing the boundary layer on the body upper surface to go through an adverse pressure gradient which could result with separated airflow. The BWB / Open Rotor configuration must be designed to mitigate these problems.

  14. Structured Overlapping Grid Simulations of Contra-rotating Open Rotor Noise

    NASA Technical Reports Server (NTRS)

    Housman, Jeffrey A.; Kiris, Cetin C.

    2015-01-01

    Computational simulations using structured overlapping grids with the Launch Ascent and Vehicle Aerodynamics (LAVA) solver framework are presented for predicting tonal noise generated by a contra-rotating open rotor (CROR) propulsion system. A coupled Computational Fluid Dynamics (CFD) and Computational AeroAcoustics (CAA) numerical approach is applied. Three-dimensional time-accurate hybrid Reynolds Averaged Navier-Stokes/Large Eddy Simulation (RANS/LES) CFD simulations are performed in the inertial frame, including dynamic moving grids, using a higher-order accurate finite difference discretization on structured overlapping grids. A higher-order accurate free-stream preserving metric discretization with discrete enforcement of the Geometric Conservation Law (GCL) on moving curvilinear grids is used to create an accurate, efficient, and stable numerical scheme. The aeroacoustic analysis is based on a permeable surface Ffowcs Williams-Hawkings (FW-H) approach, evaluated in the frequency domain. A time-step sensitivity study was performed using only the forward row of blades to determine an adequate time-step. The numerical approach is validated against existing wind tunnel measurements.

  15. Use of active control systems to improve bending and rotor flapping response of a tilt rotor VTOL airplane

    NASA Technical Reports Server (NTRS)

    Whitaker, H. P.; Cheng, Y.

    1975-01-01

    The results are summarized of an analytical study of the use of active control systems for the purpose of reducing the root mean square response of wing vertical bending and rotor flapping to atmospheric turbulence for a tilt-rotor VTOL airplane. Only the wing/rotor assembly was considered so that results of a wind tunnel test program would be applicable in a subsequent phase of the research. The capabilities and limitations of simple single feedback configurations were identified, and the most promising multiloop feedback configurations were then investigated. Design parameters were selected so as to minimize either wing bending or rotor flapping response. Within the constraints imposed by practical levels of feedback gains and complexity and by considerations of safety, reduction in response due to turbulence of the order of 30 to 50 percent is predicted using the rotor longitudinal cyclic and a trailing edge wing flap as control effectors.

  16. Measurements of the Aeroacoustic Sound Source in Hot Jets

    NASA Technical Reports Server (NTRS)

    Bridges, James; Wernet, Mark

    2004-01-01

    We have succeeded in measuring a substantial portion of the two-point space-time velocity correlation in hot, high speed turbulent jets. This measurement, crucial in aeroacoustic theory and the prediction of jet noise, has been sought for a long time, but has not been made due to the limitations of anemometry. Particle Image Velocimetry has reached a stage of maturity where sufficient measurement density in both time and space allow the computation of space-time correlations. This paper documents these measurements along with lower-order statistics to document the adherence of the jet rig and instrumentation to conventional measures of the turbulence of jets. These measures have been made for a simple round convergent nozzle at acoustic Mach numbers of 0.5, 0.9, both cold and at a static temperature ratio of 2.7, allowing some estimation of the changes in turbulence that take place with changes in jet temperature. Since the dataset described in this paper is very extensive, attention will be focused on validation of the rig and of the measurement systems, and on some of the interesting observations made from studying the statistics, especially as they relate to jet noise. Of note is the effort to study the acoustically relevant part of the space-time correlation by addressing that part of the turbulence kinetic energy that has sonic phase speed.

  17. Parallelization of an Object-Oriented Unstructured Aeroacoustics Solver

    NASA Technical Reports Server (NTRS)

    Baggag, Abdelkader; Atkins, Harold; Oezturan, Can; Keyes, David

    1999-01-01

    A computational aeroacoustics code based on the discontinuous Galerkin method is ported to several parallel platforms using MPI. The discontinuous Galerkin method is a compact high-order method that retains its accuracy and robustness on non-smooth unstructured meshes. In its semi-discrete form, the discontinuous Galerkin method can be combined with explicit time marching methods making it well suited to time accurate computations. The compact nature of the discontinuous Galerkin method also makes it well suited for distributed memory parallel platforms. The original serial code was written using an object-oriented approach and was previously optimized for cache-based machines. The port to parallel platforms was achieved simply by treating partition boundaries as a type of boundary condition. Code modifications were minimal because boundary conditions were abstractions in the original program. Scalability results are presented for the SCI Origin, IBM SP2, and clusters of SGI and Sun workstations. Slightly superlinear speedup is achieved on a fixed-size problem on the Origin, due to cache effects.

  18. Introduction to Generalized Functions with Applications in Aerodynamics and Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Farassat, F.

    1994-01-01

    Generalized functions have many applications in science and engineering. One useful aspect is that discontinuous functions can be handled as easily as continuous or differentiable functions and provide a powerful tool in formulating and solving many problems of aerodynamics and acoustics. Furthermore, generalized function theory elucidates and unifies many ad hoc mathematical approaches used by engineers and scientists. We define generalized functions as continuous linear functionals on the space of infinitely differentiable functions with compact support, then introduce the concept of generalized differentiation. Generalized differentiation is the most important concept in generalized function theory and the applications we present utilize mainly this concept. First, some results of classical analysis, are derived with the generalized function theory. Other applications of the generalized function theory in aerodynamics discussed here are the derivations of general transport theorems for deriving governing equations of fluid mechanics, the interpretation of the finite part of divergent integrals, the derivation of the Oswatitsch integral equation of transonic flow, and the analysis of velocity field discontinuities as sources of vorticity. Applications in aeroacoustics include the derivation of the Kirchhoff formula for moving surfaces, the noise from moving surfaces, and shock noise source strength based on the Ffowcs Williams-Hawkings equation.

  19. Aeroacoustic theory for noncompact wing-gust interaction

    NASA Technical Reports Server (NTRS)

    Martinez, R.; Widnall, S. E.

    1981-01-01

    Three aeroacoustic models for noncompact wing-gust interaction were developed for subsonic flow. The first is that for a two dimensional (infinite span) wing passing through an oblique gust. The unsteady pressure field was obtained by the Wiener-Hopf technique; the airfoil loading and the associated acoustic field were calculated, respectively, by allowing the field point down on the airfoil surface, or by letting it go to infinity. The second model is a simple spanwise superposition of two dimensional solutions to account for three dimensional acoustic effects of wing rotation (for a helicopter blade, or some other rotating planform) and of finiteness of wing span. A three dimensional theory for a single gust was applied to calculate the acoustic signature in closed form due to blade vortex interaction in helicopters. The third model is that of a quarter infinite plate with side edge through a gust at high subsonic speed. An approximate solution for the three dimensional loading and the associated three dimensional acoustic field in closed form was obtained. The results reflected the acoustic effect of satisfying the correct loading condition at the side edge.

  20. Automated Approach to Very High-Order Aeroacoustic Computations. Revision

    NASA Technical Reports Server (NTRS)

    Dyson, Rodger W.; Goodrich, John W.

    2001-01-01

    Computational aeroacoustics requires efficient, high-resolution simulation tools. For smooth problems, this is best accomplished with very high-order in space and time methods on small stencils. However, the complexity of highly accurate numerical methods can inhibit their practical application, especially in irregular geometries. This complexity is reduced by using a special form of Hermite divided-difference spatial interpolation on Cartesian grids, and a Cauchy-Kowalewski recursion procedure for time advancement. In addition, a stencil constraint tree reduces the complexity of interpolating grid points that am located near wall boundaries. These procedures are used to develop automatically and to implement very high-order methods (> 15) for solving the linearized Euler equations that can achieve less than one grid point per wavelength resolution away from boundaries by including spatial derivatives of the primitive variables at each grid point. The accuracy of stable surface treatments is currently limited to 11th order for grid aligned boundaries and to 2nd order for irregular boundaries.

  1. Unstructured CFD and Noise Prediction Methods for Propulsion Airframe Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Pao, S. Paul; Abdol-Hamid, Khaled S.; Campbell, Richard L.; Hunter, Craig A.; Massey, Steven J.; Elmiligui, Alaa A.

    2006-01-01

    Using unstructured mesh CFD methods for Propulsion Airframe Aeroacoustics (PAA) analysis has the distinct advantage of precise and fast computational mesh generation for complex propulsion and airframe integration arrangements that include engine inlet, exhaust nozzles, pylon, wing, flaps, and flap deployment mechanical parts. However, accurate solution values of shear layer velocity, temperature and turbulence are extremely important for evaluating the usually small noise differentials of potential applications to commercial transport aircraft propulsion integration. This paper describes a set of calibration computations for an isolated separate flow bypass ratio five engine nozzle model and the same nozzle system with a pylon. These configurations have measured data along with prior CFD solutions and noise predictions using a proven structured mesh method, which can be used for comparison to the unstructured mesh solutions obtained in this investigation. This numerical investigation utilized the TetrUSS system that includes a Navier-Stokes solver, the associated unstructured mesh generation tools, post-processing utilities, plus some recently added enhancements to the system. New features necessary for this study include the addition of two equation turbulence models to the USM3D code, an h-refinement utility to enhance mesh density in the shear mixing region, and a flow adaptive mesh redistribution method. In addition, a computational procedure was developed to optimize both solution accuracy and mesh economy. Noise predictions were completed using an unstructured mesh version of the JeT3D code.

  2. A Very High Order, Adaptable MESA Implementation for Aeroacoustic Computations

    NASA Technical Reports Server (NTRS)

    Dydson, Roger W.; Goodrich, John W.

    2000-01-01

    Since computational efficiency and wave resolution scale with accuracy, the ideal would be infinitely high accuracy for problems with widely varying wavelength scales. Currently, many of the computational aeroacoustics methods are limited to 4th order accurate Runge-Kutta methods in time which limits their resolution and efficiency. However, a new procedure for implementing the Modified Expansion Solution Approximation (MESA) schemes, based upon Hermitian divided differences, is presented which extends the effective accuracy of the MESA schemes to 57th order in space and time when using 128 bit floating point precision. This new approach has the advantages of reducing round-off error, being easy to program. and is more computationally efficient when compared to previous approaches. Its accuracy is limited only by the floating point hardware. The advantages of this new approach are demonstrated by solving the linearized Euler equations in an open bi-periodic domain. A 500th order MESA scheme can now be created in seconds, making these schemes ideally suited for the next generation of high performance 256-bit (double quadruple) or higher precision computers. This ease of creation makes it possible to adapt the algorithm to the mesh in time instead of its converse: this is ideal for resolving varying wavelength scales which occur in noise generation simulations. And finally, the sources of round-off error which effect the very high order methods are examined and remedies provided that effectively increase the accuracy of the MESA schemes while using current computer technology.

  3. Unsteady Aerodynamic Models for Turbomachinery Aeroelastic and Aeroacoustic Applications

    NASA Technical Reports Server (NTRS)

    Verdon, Joseph M.; Barnett, Mark; Ayer, Timothy C.

    1995-01-01

    Theoretical analyses and computer codes are being developed for predicting compressible unsteady inviscid and viscous flows through blade rows of axial-flow turbomachines. Such analyses are needed to determine the impact of unsteady flow phenomena on the structural durability and noise generation characteristics of the blading. The emphasis has been placed on developing analyses based on asymptotic representations of unsteady flow phenomena. Thus, high Reynolds number flows driven by small amplitude unsteady excitations have been considered. The resulting analyses should apply in many practical situations and lead to a better understanding of the relevant flow physics. In addition, they will be efficient computationally, and therefore, appropriate for use in aeroelastic and aeroacoustic design studies. Under the present effort, inviscid interaction and linearized inviscid unsteady flow models have been formulated, and inviscid and viscid prediction capabilities for subsonic steady and unsteady cascade flows have been developed. In this report, we describe the linearized inviscid unsteady analysis, LINFLO, the steady inviscid/viscid interaction analysis, SFLOW-IVI, and the unsteady viscous layer analysis, UNSVIS. These analyses are demonstrated via application to unsteady flows through compressor and turbine cascades that are excited by prescribed vortical and acoustic excitations and by prescribed blade vibrations. Recommendations are also given for the future research needed for extending and improving the foregoing asymptotic analyses, and to meet the goal of providing efficient inviscid/viscid interaction capabilities for subsonic and transonic unsteady cascade flows.

  4. Aeroelastic characteristics of the AH-64 bearingless tail rotor

    NASA Technical Reports Server (NTRS)

    Banerjee, D.

    1988-01-01

    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.

  5. A Novel Method for Reducing Rotor Blade-Vortex Interaction

    NASA Technical Reports Server (NTRS)

    Glinka, A. T.

    2000-01-01

    One of the major hindrances to expansion of the rotorcraft market is the high-amplitude noise they produce, especially during low-speed descent, where blade-vortex interactions frequently occur. In an attempt to reduce the noise levels caused by blade-vortex interactions, the flip-tip rotor blade concept was devised. The flip-tip rotor increases the miss distance between the shed vortices and the rotor blades, reducing BVI noise. The distance is increased by rotating an outboard portion of the rotor tip either up or down depending on the flight condition. The proposed plan for the grant consisted of a computational simulation of the rotor aerodynamics and its wake geometry to determine the effectiveness of the concept, coupled with a series of wind tunnel experiments exploring the value of the device and validating the computer model. The computational model did in fact show that the miss distance could be increased, giving a measure of the effectiveness of the flip-tip rotor. However, the wind experiments were not able to be conducted. Increased outside demand for the 7'x lO' wind tunnel at NASA Ames and low priority at Ames for this project forced numerous postponements of the tests, eventually pushing the tests beyond the life of the grant. A design for the rotor blades to be tested in the wind tunnel was completed and an analysis of the strength of the model blades based on predicted loads, including dynamic forces, was done.

  6. System Noise Assessment and the Potential for a Low Noise Hybrid Wing Body Aircraft with Open Rotor Propulsion

    NASA Technical Reports Server (NTRS)

    Thomas, Russell H.; Burley, Casey L.; Lopes, Leonard V.; Bahr, Christopher J.; Gern, Frank H.; VanZante, Dale E.

    2014-01-01

    An aircraft system noise assessment was conducted for a hybrid wing body freighter aircraft concept configured with three open rotor engines. The primary objective of the study was to determine the aircraft system level noise given the significant impact of installation effects including shielding the open rotor noise by the airframe. The aircraft was designed to carry a payload of 100,000 lbs on a 6,500 nautical mile mission. An experimental database was used to establish the propulsion airframe aeroacoustic installation effects including those from shielding by the airframe planform, interactions with the control surfaces, and additional noise reduction technologies. A second objective of the study applied the impacts of projected low noise airframe technology and a projection of advanced low noise rotors appropriate for the NASA N+2 2025 timeframe. With the projection of low noise rotors and installation effects, the aircraft system level was 26.0 EPNLdB below Stage 4 level with the engine installed at 1.0 rotor diameters upstream of the trailing edge. Moving the engine to 1.5 rotor diameters brought the system level noise to 30.8 EPNLdB below Stage 4. At these locations on the airframe, the integrated level of installation effects including shielding can be as much as 20 EPNLdB cumulative in addition to lower engine source noise from advanced low noise rotors. And finally, an additional set of technology effects were identified and the potential impact at the system level was estimated for noise only without assessing the impact on aircraft performance. If these additional effects were to be included it is estimated that the potential aircraft system noise could reach as low as 38.0 EPNLdB cumulative below Stage 4.

  7. A comparative analysis of XV-15 tiltrotor hover test data and WOPWOP predictions incorporating the fountain effect

    NASA Technical Reports Server (NTRS)

    Rutledge, Charles K.; Coffen, Charles D.; George, Albert R.

    1991-01-01

    Acoustic measurements from a hovering full scale XV-15 tilt rotor with the advanced technology blades are presented which show the directionality of fountain effect noise. Predicted acoustic directivity results are also presented which show agreement with the measured data. The aeroacoustic code, WOPWOP, was used in conjunction with a mathematical model which simulated the fountain recirculation aerodynamic effect on the rotor blade surface pressures. The predictions were used to identify the spike character in the measured data as fountain effect associated noise. The directivity of the fountain effect noise was observed to be dominant at the rear of the aircraft with increased intensities 45 degrees below the rotor disk planes.

  8. Unsteady Simulation of the Viscous Flow About a V-22 Rotor and Wing in Hover

    NASA Technical Reports Server (NTRS)

    Meakin, Robert L.

    1996-01-01

    Results of an unsteady thin-layer Navier-Stokes simulation of a 0.658-scale V-22 rotor and wing configuration in hover are presented. All geometric components of the flapped-wing and rotor test rig, including rotor blades, are accurately modeled. Rotor motion and rotor/airframe interference effects are simulated directly using moving body overset grid methods. Tiltrotor hover aerodynamics are visualized via unsteady particle trace images. Wing download predictive ability is demonstrated. Simulation results are compared with experimental data.

  9. A Study of the Effects of Blade Shape on Rotor Noise

    NASA Technical Reports Server (NTRS)

    Jones, Henry E.; Burley, Casey L.

    1997-01-01

    A new rotor noise prediction system called the Tiltrotor Aeroacoustic Code (TRAC) has been developed under the Short Haul (Civil Tiltrotor) program between NASA, the Army, and the U.S. helicopter industry. This system couples the comprehensive rotorcraft code CAMRAD.Mod1 with either the high resolution sectional loads code HIRES or the full potential CFD code FPRBVI to predict unsteady blade loads, which are then input to the noise prediction program WOPWOP. In this paper, HIRES will be used to predict the blade-vortex interaction (BVI) noise trends associated with blade shape. The baseline shape selected was a 17% scale model of a contemporary design 4 bladed rotor. Measurements for this rotor were acquired in the Duits-Nederslandse Windtunnel (DNW). The code is used to predict noise for the base configuration and the results are compared to the measured data. This provides a firm foundation for investigating the BVI noise trends associated with blade shape. The shapes selected for study are based on variation of sweep and taper which reflect plausible "passive" design concepts. Comparisons of power required, integrated noise, and aerodynamics are made and important trends are noted.

  10. A Resonant Synchronous Vibration Based Approach for Rotor Imbalance Detection

    NASA Technical Reports Server (NTRS)

    Luo, Huangeng; Rodriquez, Hector; Hallman, Darren; Lewicki, David G.

    2006-01-01

    This paper presents a methodology of detecting rotor imbalances, such as mass imbalance and crack-induced imbalance, using shaft synchronous vibrations. An iterative scheme is developed to identify parameters from measured synchronous vibration data. A detection system is integrated by using state-of-the-art commercial analysis equipment. A laboratory rotor test rig is used to verify the system integration and algorithm validation. A real engine test has been carried out and the results are reported.

  11. Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 4: Computer user's manual for UAAP turboprop aeroacoustic code

    NASA Technical Reports Server (NTRS)

    Menthe, R. W.; Mccolgan, C. J.; Ladden, R. M.

    1991-01-01

    The Unified AeroAcoustic Program (UAAP) code calculates the airloads on a single rotation prop-fan, or propeller, and couples these airloads with an acoustic radiation theory, to provide estimates of near-field or far-field noise levels. The steady airloads can also be used to calculate the nonuniform velocity components in the propeller wake. The airloads are calculated using a three dimensional compressible panel method which considers the effects of thin, cambered, multiple blades which may be highly swept. These airloads may be either steady or unsteady. The acoustic model uses the blade thickness distribution and the steady or unsteady aerodynamic loads to calculate the acoustic radiation. The users manual for the UAAP code is divided into five sections: general code description; input description; output description; system description; and error codes. The user must have access to IMSL10 libraries (MATH and SFUN) for numerous calls made for Bessel functions and matrix inversion. For plotted output users must modify the dummy calls to plotting routines included in the code to system-specific calls appropriate to the user's installation.

  12. Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 4: Computer user's manual for UAAP turboprop aeroacoustic code

    NASA Astrophysics Data System (ADS)

    Menthe, R. W.; McColgan, C. J.; Ladden, R. M.

    1991-05-01

    The Unified AeroAcoustic Program (UAAP) code calculates the airloads on a single rotation prop-fan, or propeller, and couples these airloads with an acoustic radiation theory, to provide estimates of near-field or far-field noise levels. The steady airloads can also be used to calculate the nonuniform velocity components in the propeller wake. The airloads are calculated using a three dimensional compressible panel method which considers the effects of thin, cambered, multiple blades which may be highly swept. These airloads may be either steady or unsteady. The acoustic model uses the blade thickness distribution and the steady or unsteady aerodynamic loads to calculate the acoustic radiation. The users manual for the UAAP code is divided into five sections: general code description; input description; output description; system description; and error codes. The user must have access to IMSL10 libraries (MATH and SFUN) for numerous calls made for Bessel functions and matrix inversion. For plotted output users must modify the dummy calls to plotting routines included in the code to system-specific calls appropriate to the user's installation.

  13. Vibration sensing in smart machine rotors using internal MEMS accelerometers

    NASA Astrophysics Data System (ADS)

    Jiménez, Samuel; Cole, Matthew O. T.; Keogh, Patrick S.

    2016-09-01

    This paper presents a novel topology for enhanced vibration sensing in which wireless MEMS accelerometers embedded within a hollow rotor measure vibration in a synchronously rotating frame of reference. Theoretical relations between rotor-embedded accelerometer signals and the vibration of the rotor in an inertial reference frame are derived. It is thereby shown that functionality as a virtual stator-mounted displacement transducer can be achieved through appropriate signal processing. Experimental tests on a prototype rotor confirm that both magnitude and phase information of synchronous vibration can be measured directly without additional stator-mounted key-phasor sensors. Displacement amplitudes calculated from accelerometer signals will become erroneous at low rotational speeds due to accelerometer zero-g offsets, hence a corrective procedure is introduced. Impact tests are also undertaken to examine the ability of the internal accelerometers to measure transient vibration. A further capability is demonstrated, whereby the accelerometer signals are used to measure rotational speed of the rotor by analysing the signal component due to gravity. The study highlights the extended functionality afforded by internal accelerometers and demonstrates the feasibility of internal sensor topologies, which can provide improved observability of rotor vibration at externally inaccessible rotor locations.

  14. Traversing Microphone Track Installed in NASA Lewis' Aero-Acoustic Propulsion Laboratory Dome

    NASA Technical Reports Server (NTRS)

    Bauman, Steven W.; Perusek, Gail P.

    1999-01-01

    The Aero-Acoustic Propulsion Laboratory is an acoustically treated, 65-ft-tall dome located at the NASA Lewis Research Center. Inside this laboratory is the Nozzle Acoustic Test Rig (NATR), which is used in support of Advanced Subsonics Technology (AST) and High Speed Research (HSR) to test engine exhaust nozzles for thrust and acoustic performance under simulated takeoff conditions. Acoustic measurements had been gathered by a far-field array of microphones located along the dome wall and 10-ft above the floor. Recently, it became desirable to collect acoustic data for engine certifications (as specified by the Federal Aviation Administration (FAA)) that would simulate the noise of an aircraft taking off as heard from an offset ground location. Since nozzles for the High-Speed Civil Transport have straight sides that cause their noise signature to vary radially, an additional plane of acoustic measurement was required. Desired was an arched array of 24 microphones, equally spaced from the nozzle and each other, in a 25 off-vertical plane. The various research requirements made this a challenging task. The microphones needed to be aimed at the nozzle accurately and held firmly in place during testing, but it was also essential that they be easily and routinely lowered to the floor for calibration and servicing. Once serviced, the microphones would have to be returned to their previous location near the ceiling. In addition, there could be no structure could between the microphones and the nozzle, and any structure near the microphones would have to be designed to minimize noise reflections. After many concepts were considered, a single arched truss structure was selected that would be permanently affixed to the dome ceiling and to one end of the dome floor.

  15. Sound Emission of Rotor Induced Deformations of Generator Casings

    NASA Technical Reports Server (NTRS)

    Polifke, W.; Mueller, B.; Yee, H. C.; Mansour, Nagi (Technical Monitor)

    2001-01-01

    The casing of large electrical generators can be deformed slightly by the rotor's magnetic field. The sound emission produced by these periodic deformations, which could possibly exceed guaranteed noise emission limits, is analysed analytically and numerically. From the deformation of the casing, the normal velocity of the generator's surface is computed. Taking into account the corresponding symmetry, an analytical solution for the acoustic pressure outside the generator is round in terms of the Hankel function of second order. The normal velocity or the generator surface provides the required boundary condition for the acoustic pressure and determines the magnitude of pressure oscillations. For the numerical simulation, the nonlinear 2D Euler equations are formulated In a perturbation form for low Mach number Computational Aeroacoustics (CAA). The spatial derivatives are discretized by the classical sixth-order central interior scheme and a third-order boundary scheme. Spurious high frequency oscillations are damped by a characteristic-based artificial compression method (ACM) filter. The time derivatives are approximated by the classical 4th-order Runge-Kutta method. The numerical results are In excellent agreement with the analytical solution.

  16. Variable camber rotor study

    NASA Technical Reports Server (NTRS)

    Dadone, L.; Cowan, J.; Mchugh, F. J.

    1982-01-01

    Deployment of variable camber concepts on helicopter rotors was analytically assessed. It was determined that variable camber extended the operating range of helicopters provided that the correct compromise can be obtained between performance/loads gains and mechanical complexity. A number of variable camber concepts were reviewed on a two dimensional basis to determine the usefulness of leading edge, trailing edge and overall camber variation schemes. The most powerful method to vary camber was through the trailing edge flaps undergoing relatively small motions (-5 deg to +15 deg). The aerodynamic characteristics of the NASA/Ames A-1 airfoil with 35% and 50% plain trailing edge flaps were determined by means of current subcritical and transonic airfoil design methods and used by rotor performance and loads analysis codes. The most promising variable camber schedule reviewed was a configuration with a 35% plain flap deployment in an on/off mode near the tip of a blade. Preliminary results show approximately 11% reduction in power is possible at 192 knots and a rotor thrust coefficient of 0.09. The potential demonstrated indicates a significant potential for expanding the operating envelope of the helicopter. Further investigation into improving the power saving and defining the improvement in the operational envelope of the helicopter is recommended.

  17. Polygonal shaft hole rotor

    DOEpatents

    Hussey, John H.; Rose, John Scott; Meystrik, Jeffrey J.; White, Kent Lee

    2001-01-23

    A laminated rotor for an induction motor has a plurality of ferro-magnetic laminations mounted axially on a rotor shaft. Each of the plurality of laminations has a central aperture in the shape of a polygon with sides of equal length. The laminations are alternatingly rotated 180.degree. from one another so that the straight sides of the polygon shaped apertures are misaligned. As a circular rotor shaft is press fit into a stack of laminations, the point of maximum interference occurs at the midpoints of the sides of the polygon (i.e., at the smallest radius of the central apertures of the laminations). Because the laminates are alternatingly rotated, the laminate material at the points of maximum interference yields relatively easily into the vertices (i.e., the greatest radius of the central aperture) of the polygonal central aperture of the next lamination as the shaft is inserted into the stack of laminations. Because of this yielding process, the amount of force required to insert the shaft is reduced, and a tighter fit is achieved.

  18. Single rotor turbine engine

    DOEpatents

    Platts, David A.

    2002-01-01

    There has been invented a turbine engine with a single rotor which cools the engine, functions as a radial compressor, pushes air through the engine to the ignition point, and acts as an axial turbine for powering the compressor. The invention engine is designed to use a simple scheme of conventional passage shapes to provide both a radial and axial flow pattern through the single rotor, thereby allowing the radial intake air flow to cool the turbine blades and turbine exhaust gases in an axial flow to be used for energy transfer. In an alternative embodiment, an electric generator is incorporated in the engine to specifically adapt the invention for power generation. Magnets are embedded in the exhaust face of the single rotor proximate to a ring of stationary magnetic cores with windings to provide for the generation of electricity. In this alternative embodiment, the turbine is a radial inflow turbine rather than an axial turbine as used in the first embodiment. Radial inflow passages of conventional design are interleaved with radial compressor passages to allow the intake air to cool the turbine blades.

  19. Dynamic Tester For Rotor Seals And Bearings

    NASA Technical Reports Server (NTRS)

    Von Pragenau, George L.

    1991-01-01

    Simplified apparatus measures performance under vibration. Measures some of dynamic parameters of rotor seals and bearings. Tests damping seals, damping bearings, conventional seals, and conventional bearings. Used with variety of pumped liquids, from water to liquid oxygen or hydrogen. Designed to test bearings and seals of turbopumps, tester rotates shaft at high speed while liquid flows much as it would in real turbopump. Also measures deflections of components.

  20. Multicyclic Controllable Twist Rotor Data Analysis

    NASA Technical Reports Server (NTRS)

    Wei, F. S.; Weisbrich, A. L.

    1979-01-01

    Rsults provide functional relationship between rotor performance, blade vibratory loads and dual control settings and indicate that multicyclic control produced significant reductions in blade flatwise bending moments and blade root actuator control loads. Higher harmonic terms of servo flap deflection were found to be most pronounced in flatwise bending moment, transmission vertical vibration and pitch link vibratory load equations. The existing test hardware represents a satisfactory configuration for demonstrating MCTR technology and defining a data base for additional wind tunnel testing.

  1. Navier-Stokes analysis of radial turbine rotor performance

    NASA Technical Reports Server (NTRS)

    Larosiliere, L. M.

    1993-01-01

    An analysis of flow through a radial turbine rotor using the three-dimensional, thin-layer Navier-Stokes code RVC3D is described. The rotor is a solid version of an air-cooled metallic radial turbine having thick trailing edges, shroud clearance, and scalloped-backface clearance. Results are presented at the nominal operating condition using both a zero-clearance model and a model simulating the effects of the shroud and scalloped-backface clearance flows. A comparison with the available test data is made and details of the internal flow physics are discussed, allowing a better understanding of the complex flow distribution within the rotor.

  2. Efficient Helicopter Aerodynamic and Aeroacoustic Predictions on Parallel Computers

    NASA Technical Reports Server (NTRS)

    Wissink, Andrew M.; Lyrintzis, Anastasios S.; Strawn, Roger C.; Oliker, Leonid; Biswas, Rupak

    1996-01-01

    This paper presents parallel implementations of two codes used in a combined CFD/Kirchhoff methodology to predict the aerodynamics and aeroacoustics properties of helicopters. The rotorcraft Navier-Stokes code, TURNS, computes the aerodynamic flowfield near the helicopter blades and the Kirchhoff acoustics code computes the noise in the far field, using the TURNS solution as input. The overall parallel strategy adds MPI message passing calls to the existing serial codes to allow for communication between processors. As a result, the total code modifications required for parallel execution are relatively small. The biggest bottleneck in running the TURNS code in parallel comes from the LU-SGS algorithm that solves the implicit system of equations. We use a new hybrid domain decomposition implementation of LU-SGS to obtain good parallel performance on the SP-2. TURNS demonstrates excellent parallel speedups for quasi-steady and unsteady three-dimensional calculations of a helicopter blade in forward flight. The execution rate attained by the code on 114 processors is six times faster than the same cases run on one processor of the Cray C-90. The parallel Kirchhoff code also shows excellent parallel speedups and fast execution rates. As a performance demonstration, unsteady acoustic pressures are computed at 1886 far-field observer locations for a sample acoustics problem. The calculation requires over two hundred hours of CPU time on one C-90 processor but takes only a few hours on 80 processors of the SP2. The resultant far-field acoustic field is analyzed with state of-the-art audio and video rendering of the propagating acoustic signals.

  3. Study of rotor gust response by means of the local momentum theory

    NASA Technical Reports Server (NTRS)

    Azuma, A.; Saito, S.

    1982-01-01

    The vertical gust response of the helicopter rotor in cruising flight is studied analytically by means of the local momentum theory (LMT) and experimentally in a wind tunnel utilizing a gust generator. By introducing the unsteady aerodynamic effects into the LMT and by considering the elastic deformation of the rotor blade, the vibratory characteristics of flapping blades and of the rotor forces can be obtained. Since the LMT makes it possible to calculate the instantaneous load distribution along the rotor blade in desired azimuthal or timewise intervals, the effects of gradual penetration of the rotor into the gust can be studied and the Fourier analysis or power spectrum analysis of the rotor response applied to any kind of gust input. Some results obtained analytically are verified by the experimental tests performed in the wind tunnel which generates step, sinusoidal and random vertical gusts by the motion of a series of cascaded vanes upstream of the rotor.

  4. Acoustic predictions using measured pressures from a model rotor in the DNW

    NASA Technical Reports Server (NTRS)

    Visintainer, Joseph A.; Burley, Casey L.; Marcolini, Michael A.; Liu, Sandy R.

    1991-01-01

    A contemporary design, 4-bladed United Technologies model rotor with pressure-instrumented blades was tested in the Duits-Nederslandse Windtunnel. Simultaneous acoustic and pressure measurements were made for a wide range of operating conditions. Microphones were optimally positioned at a number of locations in the flow forward of the rotor to measure rotor thickness noise, high-speed impulsive noise (both in the rotor plane), and blade-vortex interaction noise (forward and 25 deg below the rotor plane). The blade surface pressure data are used as aerodynamic input to WOPWOP, which is a state-of-the-art rotor noise prediction program that predicts rotor thickness and loading noise. The predicted results using WOPWOP are compared to the measured noise levels for cases where either thickness noise, blade-vortex interaction noise, or high-speed impulsive noise is the dominant noise mechanism. The comparisons show regions of good agreement, as well as areas where further improvement is necessary.

  5. The use of staggered scheme and an absorbing buffer zone for computational aeroacoustics

    NASA Technical Reports Server (NTRS)

    Nark, Douglas M.

    1995-01-01

    Various problems from those proposed for the Computational Aeroacoustics (CAA) workshop were studied using second and fourth order staggered spatial discretizations in conjunction with fourth order Runge-Kutta time integration. In addition, an absorbing buffer zone was used at the outflow boundaries. Promising results were obtained and provide a basis for application of these techniques to a wider variety of problems.

  6. Application of NASA General-Purpose Solver to Large-Scale Computations in Aeroacoustics

    NASA Technical Reports Server (NTRS)

    Watson, Willie R.; Storaasli, Olaf O.

    2004-01-01

    Of several iterative and direct equation solvers evaluated previously for computations in aeroacoustics, the most promising was the NASA-developed General-Purpose Solver (winner of NASA's 1999 software of the year award). This paper presents detailed, single-processor statistics of the performance of this solver, which has been tailored and optimized for large-scale aeroacoustic computations. The statistics, compiled using an SGI ORIGIN 2000 computer with 12 Gb available memory (RAM) and eight available processors, are the central processing unit time, RAM requirements, and solution error. The equation solver is capable of solving 10 thousand complex unknowns in as little as 0.01 sec using 0.02 Gb RAM, and 8.4 million complex unknowns in slightly less than 3 hours using all 12 Gb. This latter solution is the largest aeroacoustics problem solved to date with this technique. The study was unable to detect any noticeable error in the solution, since noise levels predicted from these solution vectors are in excellent agreement with the noise levels computed from the exact solution. The equation solver provides a means for obtaining numerical solutions to aeroacoustics problems in three dimensions.

  7. Homopolar motor with dual rotors

    DOEpatents

    Hsu, J.S.

    1998-12-01

    A homopolar motor has a field rotor mounted on a frame for rotation in a first rotational direction and for producing an electromagnetic field, and an armature rotor mounted for rotation on said frame within said electromagnetic field and in a second rotational direction counter to said first rotational direction of said field rotor. The two rotors are coupled through a 1:1 gearing mechanism, so as to travel at the same speed but in opposite directions. This doubles the output voltage and output power, as compared to a motor in which only the armature is rotated. Several embodiments are disclosed. 7 figs.

  8. Homopolar motor with dual rotors

    DOEpatents

    Hsu, John S.

    1998-01-01

    A homopolar motor (10) has a field rotor (15) mounted on a frame (11) for rotation in a first rotational direction and for producing an electromagnetic field, and an armature rotor (17) mounted for rotation on said frame (11) within said electromagnetic field and in a second rotational direction counter to said first rotational direction of said field rotor (15). The two rotors (15, 17) are coupled through a 1:1 gearing mechanism (19), so as to travel at the same speed but in opposite directions. This doubles the output voltage and output power, as compared to a motor in which only the armature is rotated. Several embodiments are disclosed.

  9. 14 CFR 29.921 - Rotor brake.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Rotor Drive System § 29.921 Rotor brake. If there is a means to control the rotation of the rotor drive system independently of the engine, any limitations...

  10. 14 CFR 27.921 - Rotor brake.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... STANDARDS: NORMAL CATEGORY ROTORCRAFT Powerplant Rotor Drive System § 27.921 Rotor brake. If there is a means to control the rotation of the rotor drive system independently of the engine, any limitations...

  11. Assessment of Scaled Rotors for Wind Tunnel Experiments.

    SciTech Connect

    Maniaci, David Charles; Kelley, Christopher Lee; Chiu, Phillip

    2015-07-01

    Rotor design and analysis work has been performed to support the conceptualization of a wind tunnel test focused on studying wake dynamics. This wind tunnel test would serve as part of a larger model validation campaign that is part of the Department of Energy Wind and Water Power Program’s Atmosphere to electrons (A2e) initiative. The first phase of this effort was directed towards designing a functionally scaled rotor based on the same design process and target full-scale turbine used for new rotors for the DOE/SNL SWiFT site. The second phase focused on assessing the capabilities of an already available rotor, the G1, designed and built by researchers at the Technical University of München.

  12. Numerical and Experimental Simulation of a Vertical High Speed Motorcompressor Rotor Drop onto Catcher Bearings

    NASA Astrophysics Data System (ADS)

    Ransom, David; Masala, Andrea; Moore, Jeffrey; Vannini, Giuseppe; Camatti, Massimo

    A new research program was jointly set up between GE Oil&Gas and Southwest Research Institute (SwRI), to predict and test the dynamics of a vertical rotor drop on catcher bearings. A numerical tool able to account for flexible rotor and stator dynamics, catcher bearing stiffness and damping mechanism was developed. An experimental activity on a new vertical rotor test rig was carried out. A first analysis of numerical simulations and experimental analysis is presented in this paper.

  13. Integrated technology rotor/flight research rotor concept definition study

    NASA Technical Reports Server (NTRS)

    Carlson, R. G.; Beno, E. A.; Ulisnik, H. D.

    1983-01-01

    As part of the Integrated Technology Rotor/Flight Research Rotor (ITR/FRR) Program a number of advanced rotor system designs were conceived and investigated. From these, several were chosen that best meet the started ITR goals with emphasis on stability, reduced weight and hub drag, simplicity, low head moment stiffness, and adequate strength and fatigue life. It was concluded that obtaining low hub moment stiffness was difficult when only the blade flexibility of bearingless rotor blades is considered, unacceptably low fatigue life being the primary problem. Achieving a moderate hub moment stiffness somewhat higher than state of the art articulated rotors in production today is possible within the fatigue life constraint. Alternatively, low stiffness is possible when additional rotor elements, besides the blades themselves, provide part of the rotor flexibility. Two primary designs evolved as best meeting the general ITR requirements that presently exist. An I shaped flexbeam with an external torque tube can satisfy the general goals but would have either higher stiffness or reduced fatigue life. The elastic gimbal rotor can achieve a better combination of low stiffness and high fatigue life but would be a somewhat heavier design and possibly exhibit a higher risk of aeromechanical instability.

  14. Pressure Sensitive Paint Measurements on 15% Scale Rotor Blades in Hover

    NASA Technical Reports Server (NTRS)

    Wong, Oliver D.; Watkins, Anthony Neal; Ingram, JoAnne L.

    2005-01-01

    This paper describes a proof of concept test to examine the feasibility of using pressure sensitive paint (PSP) to measure the pressure distributions on a rotor in hover. The test apparatus consisted of the US Army 2-meter Rotor Test Stand (2MRTS) and 15% scale swept tip rotor blades. Two camera/rotor separations were examined: 0.76 and 1.35 radii. The outer 15% of each blade was painted with PSP. Intensity and lifetime based PSP measurement techniques were attempted. Data were collected from all blades at thrust coefficients ranging from 0.004 to 0.009.

  15. Influence of parameter changes to stability behavior of rotors

    NASA Technical Reports Server (NTRS)

    Fritzen, C. P.; Nordmann, R.

    1982-01-01

    The occurrence of unstable vibrations in rotating machinery requires corrective measures for improvement of the stability behavior. A simple approximate method is represented to find out the influence of parameter changes to the stability behavior. The method is based on an expansion of the eigenvalues in terms of system parameters. Influence coefficients show the effect of structural modifications. The method first of all was applied to simple nonconservative rotor models. It was approved for an unsymmetric rotor of a test rig.

  16. An experimental and analytical method for approximate determination of the tilt rotor research aircraft rotor/wing download

    NASA Technical Reports Server (NTRS)

    Jordon, D. E.; Patterson, W.; Sandlin, D. R.

    1985-01-01

    The XV-15 Tilt Rotor Research Aircraft download phenomenon was analyzed. This phenomenon is a direct result of the two rotor wakes impinging on the wing upper surface when the aircraft is in the hover configuration. For this study the analysis proceeded along tow lines. First was a method whereby results from actual hover tests of the XV-15 aircraft were combined with drag coefficient results from wind tunnel tests of a wing that was representative of the aircraft wing. Second, an analytical method was used that modeled that airflow caused gy the two rotors. Formulas were developed in such a way that acomputer program could be used to calculate the axial velocities were then used in conjunction with the aforementioned wind tunnel drag coefficinet results to produce download values. An attempt was made to validate the analytical results by modeling a model rotor system for which direct download values were determinrd..

  17. Simulation of rotor blade element turbulence

    NASA Technical Reports Server (NTRS)

    Mcfarland, R. E.; Duisenberg, Ken

    1995-01-01

    A piloted, motion-based simulation of Sikorsky's Black Hawk helicopter was used as a platform for the investigation of rotorcraft responses to vertical turbulence. By using an innovative temporal and geometrical distribution algorithm that preserved the statistical characteristics of the turbulence over the rotor disc, stochastic velocity components were applied at each of twenty blade-element stations. This model was implemented on NASA Ames' Vertical Motion Simulator (VMS), and ten test pilots were used to establish that the model created realistic cues. The objectives of this research included the establishment of a simulation-technology basis for future investigation into real-time turbulence modeling. This goal was achieved; our extensive additions to the rotor model added less than a 10 percent computational overhead. Using a VAX 9000 computer the entire simulation required a cycle time of less than 12 msec. Pilot opinion during this simulation was generally quite favorable. For low speed flight the consensus was that SORBET (acronym for title) was better than the conventional body-fixed model, which was used for comparison purposes, and was determined to be too violent (like a washboard). For high speed flight the pilots could not identify differences between these models. These opinions were something of a surprise because only the vertical turbulence component on the rotor system was implemented in SORBET. Because of the finite-element distribution of the inputs, induced outputs were observed in all translational and rotational axes. Extensive post-simulation spectral analyses of the SORBET model suggest that proper rotorcraft turbulence modeling requires that vertical atmospheric disturbances not be superimposed at the vehicle center of gravity but, rather, be input into the rotor system, where the rotor-to-body transfer function severely attenuates high frequency rotorcraft responses.

  18. Development of Mach scale rotors with composite tailored couplings for vibration reduction

    NASA Astrophysics Data System (ADS)

    Bao, Jinsong

    The use of composite tailored couplings in rotor blades to reduce vibratory hub loads was studied through design, structural and aeroelastic analysis, fabrication, and wind tunnel test of Mach scale articulated composite rotors with tailored flap-bending/torsion couplings. The rotor design was nominally based on the UH-60 BLACK HAWK rotor. The 6-foot diameter blades have a SC1095 profile and feature a linear twist of -12 deg. The analysis of composite rotor was carried out using a mixed cross-section structural model, and UMARC. Five sets of composite rotor were fabricated, including a baseline rotor without coupling, rotors with spanwise uniform positive coupling and negative coupling, and rotors with spanwise dual-segmented coupling (FBT-P/N) and triple-segmented coupling. The blade composite D-spar is the primary structural element supporting the blade loads and providing the desired elastic couplings. Non-rotating tests were performed to examine blade structural properties. The measurements showed good correlation with predictions, and good repeatability for the four blades of each rotor set. All rotors were tested at a rotor speed of 2300 rpm (tip Mach number 0.65) at different advance ratios and thrust levels, in the Glenn L. Martin Wind Tunnel at the University of Maryland. The test results showed that flap-bending/torsion couplings have a significant effect on the rotor vibratory hub loads. All coupled rotors reduced the 4/rev vertical force for advance ratios up to 0.3, with reductions ranging from 1 to 34%. The mixed coupling rotor FBT-P/N reduced overall 4/rev hub loads at advance ratios of 0.1, 0.2 and 0.3. At a rotor speed of 2300 rpm and an advance ratio of 0.3, the FBT-P/N rotor achieved 15% reduction for 4/rev vertical force, 3% for 4/rev in-plane force and 14% for 4/rev head moment. The reductions in the 4/rev hub loads are related to the experimentally observed reductions in 3/rev and 5/rev blade flap bending moments. Through the present research

  19. Three-Dimensional Application of DAMAS Methodology for Aeroacoustic Noise Source Definition

    NASA Technical Reports Server (NTRS)

    Brooks, Thomas F.; Humphreys, William M., Jr.

    2005-01-01

    At the 2004 AIAA/CEAS Aeroacoustic Conference, a breakthrough in acoustic microphone array technology was reported by the authors. A Deconvolution Approach for the Mapping of Acoustic Sources (DAMAS) was developed which decouples the array design and processing influence from the noise being measured, using a simple and robust algorithm. For several prior airframe noise studies, it was shown to permit an unambiguous and accurate determination of acoustic source position and strength. As a follow-on effort, this paper examines the technique for three-dimensional (3D) applications. First, the beamforming ability for arrays, of different size and design, to focus longitudinally and laterally is examined for a range of source positions and frequency. Advantage is found for larger array designs with higher density microphone distributions towards the center. After defining a 3D grid generalized with respect to the array s beamforming characteristics, DAMAS is employed in simulated and experimental noise test cases. It is found that spatial resolution is much less sharp in the longitudinal direction in front of the array compared to side-to-side lateral resolution. 3D DAMAS becomes useful for sufficiently large arrays at sufficiently high frequency. But, such can be a challenge to computational capabilities, with regard to the required expanse and number of grid points. Also, larger arrays can strain basic physical modeling assumptions that DAMAS and all traditional array methodologies use. An important experimental result is that turbulent shear layers can negatively impact attainable beamforming resolution. Still, the usefulness of 3D DAMAS is demonstrated by the measurement of landing gear noise source distributions in a difficult hard-wall wind tunnel environment.

  20. Reducing Propulsion Airframe Aeroacoustic Interactions with Uniquely Tailored Chevrons. 1.; Isolated Nozzles

    NASA Technical Reports Server (NTRS)

    Mengle, Vinod G.; Elkroby, Ronen; Brunsniak, Leon; Thomas, Russ H.

    2006-01-01

    The flow/acoustic environment surrounding an engine nozzle installed on an airplane, say, under the wing, is asymmetric due to the pylon, the wing and the interaction of the exhaust jet with flaps on the wing. However, the conventional chevrons, which are azimuthally uniform serrations on the nozzle lip, do not exploit the asymmetry due to these propulsion airframe aeroacoustic interactions to reduce jet noise. In this pioneering study we use this non-axisymmetry to our advantage and examine if the total jet-related noise radiated to the ground can be reduced by using different types of azimuthally varying chevrons (AVC) which vary the mixing around the nozzle periphery. Several scale models of the isolated nozzle, representative of high bypass ratio engine nozzles, were made with a pylon and azimuthally varying chevrons on both fan and core nozzles to enhance mixing at the top (near the pylon) with less mixing at the bottom (away from the pylon) or vice versa. Various combinations of fan and core AVC nozzles were systematically tested at typical take-off conditions inside a free jet wind-tunnel and, here, in Part 1 we analyze the acoustics results for the isolated nozzle with a pylon, with installation effects reported in Parts 2 and 3. Several interesting results are discovered: amongst the fan AVCs the top-enhanced mixing T-fan chevron nozzle is quieter in combination with any core AVC nozzle when compared to conventional chevrons; however, the bottom-mixing B-fan chevrons, as well as the core AVC nozzles, by themselves, are noisier. Further, the low-frequency source strengths in the jet plume, obtained via phased microphone arrays, also corroborate the far field sound, and for the T-fan chevrons such sources move further downstream than those for baseline or conventional chevron nozzles.