Rotor hubs are predominantly responsible for the parasitic drag encountered by high-speed rotorcraft. To gain a deeper understanding of the fluid dynamics around rotor hubs, simulations of counter-rotating coaxial rotor-hub flows were conducted, which was subsequently followed by a corroborative experiment at Pennsylvania State University. The simulation process employed a computational fluid dynamics framework termed “Mercury,” which utilizes an unstructured⁄Cartesian multimesh paradigm. This process incorporated Spalart–Allmaras delayed detached eddy simulation turbulence modeling for an accurate representation of the flow. The investigation into the coaxial hub flow physics was conducted at two different advance ratios: 0.25 and 0.6, constructed by building up the hub components. The interaction between the hub and fairing components led to an increase in the average hub drag and caused unsteady harmonics in the hub and fairing drags. Furthermore, it was noted that different advance ratios affected the drag and wake structures. The complete hub model was then simulated in a water tunnel with an advance ratio of 0.25. Predictions of mean and unsteady drag, as well as mean wake velocity fields, were compared with the experimental results. Overall, the mean wake velocity field from the simulation qualitatively aligned with the experimental results, especially in the near-wake region. Additionally, the buildup model analysis significantly aided in understanding the intricate wake structures surrounding the hub.
{"title":"Prediction of Coaxial Rotor Hub Flow Using Mercury Framework","authors":"Yong Su Jung, Bumseok Lee, James Baeder","doi":"10.4050/jahs.69.022004","DOIUrl":"https://doi.org/10.4050/jahs.69.022004","url":null,"abstract":"Rotor hubs are predominantly responsible for the parasitic drag encountered by high-speed rotorcraft. To gain a deeper understanding of the fluid dynamics around rotor hubs, simulations of counter-rotating coaxial rotor-hub flows were conducted, which was subsequently followed by a corroborative experiment at Pennsylvania State University. The simulation process employed a computational fluid dynamics framework termed “Mercury,” which utilizes an unstructured⁄Cartesian multimesh paradigm. This process incorporated Spalart–Allmaras delayed detached eddy simulation turbulence modeling for an accurate representation of the flow. The investigation into the coaxial hub flow physics was conducted at two different advance ratios: 0.25 and 0.6, constructed by building up the hub components. The interaction between the hub and fairing components led to an increase in the average hub drag and caused unsteady harmonics in the hub and fairing drags. Furthermore, it was noted that different advance ratios affected the drag and wake structures. The complete hub model was then simulated in a water tunnel with an advance ratio of 0.25. Predictions of mean and unsteady drag, as well as mean wake velocity fields, were compared with the experimental results. Overall, the mean wake velocity field from the simulation qualitatively aligned with the experimental results, especially in the near-wake region. Additionally, the buildup model analysis significantly aided in understanding the intricate wake structures surrounding the hub.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135314406","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This study examines the acoustic behavior in hover of manned -size, multirotor, eVTOL aircraft in the classical quadcopter, hexacopter, and octocopter configurations. The rotors are assumed to have collective pitch control and operate at a specified RPM, with orthogonal and tip-to-tip rotor phasing considered. All configurations have the same disk loading and tip Mach number, with the rotor radius decreasing and RPM increasing, going from the quadcopter to the octocopter. The simulations use the Rensselaer Multicopter Analysis Code for the aerodynamic loads on the blades, coupled to an acoustic propagation code for noise predictions at selected observer locations. From the simulation results, orthogonal phasing between rotors is shown to produce significant noise reductions along interboom bisectors (between 9 and 14 dB relative to an equivalent single rotor, at 6 lb/ft2 disk loading and 0.51 tip Mach number). Further reducing the tip Mach number not only reduces the propagated noise but produces even deeper regions of quiet along the interboom bisectors (18–25 dB quieter at 3 lb/ft2 with 0.36 tip Mach number). An examination of the sound pressure level frequency spectra indicates that smaller faster spinning rotors (going from the quadcopter to octocopter) produce more tonal peaks at higher frequencies which would result in penalties in A-weighted noise.
{"title":"A Comparison of Tonal Noise Characteristics of Large Multicopters with Phased Rotors","authors":"Brendan Smith, F. Gandhi, R. Niemiec","doi":"10.4050/jahs.68.032008","DOIUrl":"https://doi.org/10.4050/jahs.68.032008","url":null,"abstract":"This study examines the acoustic behavior in hover of manned -size, multirotor, eVTOL aircraft in the classical quadcopter, hexacopter, and octocopter configurations. The rotors are assumed to have collective pitch control and operate at a specified RPM, with orthogonal and tip-to-tip rotor phasing considered. All configurations have the same disk loading and tip Mach number, with the rotor radius decreasing and RPM increasing, going from the quadcopter to the octocopter. The simulations use the Rensselaer Multicopter Analysis Code for the aerodynamic loads on the blades, coupled to an acoustic propagation code for noise predictions at selected observer locations. From the simulation results, orthogonal phasing between rotors is shown to produce significant noise reductions along interboom bisectors (between 9 and 14 dB relative to an equivalent single rotor, at 6 lb/ft2 disk loading and 0.51 tip Mach number). Further reducing the tip Mach number not only reduces the propagated noise but produces even deeper regions of quiet along the interboom bisectors (18–25 dB quieter at 3 lb/ft2 with 0.36 tip Mach number). An examination of the sound pressure level frequency spectra indicates that smaller faster spinning rotors (going from the quadcopter to octocopter) produce more tonal peaks at higher frequencies which would result in penalties in A-weighted noise.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70219522","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
David A. Coleman, Atanu Halder, Farid Saemi, Carl Runco, H. Denton, Bochan Lee, Vishaal Subramanian, Eric Greenwood, Vinod Lakshminaryan, Moble Benedict
This paper describes the development and flight testing of a personal air vehicle by team Harmony for the GoFly Prize competition. The competition emphasized an open‐air flight experience, and the final aircraft was scored by size (≤8.5 ft), noise (<87 dBA), and speed (>30 kt). The highest scoring team would win a $1 million grand prize. Team Harmony endeavored to develop an aircraft which would maximize the competition score. A counterrotating coaxial electric helicopter configuration was chosen to maximize rotor area and reduce disk loading for efficiency and acoustic benefits. The rotors were designed through a parametric study using an in‐house aerodynamics code, an in‐house acoustics code, and CREATE™‐AV Helios CFD. A quiet electric power train and custom 11 kWh battery were developed. Flight control was implemented with dual, independent, electronically coupled swashplates. A one‐third‐scale prototype aircraft was first developed and flight‐tested. Then a full‐scale, 520 lb (235.4 kg) prototype with an 8.45 ft (2.58 m) rotor diameter was developed and flight‐tested. During hovering, the measured sound pressure levels at 50 ft (15.24 m) were 73 dBA. The full‐scale vehicle crashed before its speed capabilities could be tested, but the aircraft would have likely scored well, achieving the goal of the project. This work encompassed several design trade‐offs, especially between aerodynamic and acoustic performance in the rotor and blade design, and between acoustics and endurance in the power source selection, and in availability and flight‐worthiness of off‐the‐shelf hardware for personal flight.
本文介绍了Harmony团队为GoFly奖竞赛开发的一种个人飞行器的飞行测试。比赛强调露天飞行体验，最终的飞机按尺寸(≤8.5英尺)和噪音(30千瓦)进行评分。得分最高的团队将赢得100万美元的大奖。和谐队努力研制一种能使比赛得分最大化的飞机。选择了一种反向旋转同轴电动直升机结构，以最大化旋翼面积并减少磁盘负载，从而提高效率和声学效益。旋翼的设计采用了内部空气动力学代码、内部声学代码和CREATE™- AV Helios CFD进行参数化研究。一个安静的电力传动系统和定制11千瓦时的电池被开发。飞行控制采用双、独立、电子耦合斜盘。一架三分之一比例的原型飞机首先被开发出来并进行了飞行测试。然后，一个全尺寸，520磅(235.4公斤)原型机与8.45英尺(2.58米)转子直径被开发和飞行测试。悬停时，在50英尺(15.24米)处测量到的声压级为73 dBA。在测试其速度能力之前，全尺寸飞行器坠毁了，但飞机可能会取得很好的成绩，实现了项目的目标。这项工作包括几个设计权衡，特别是在转子和叶片设计中的空气动力学和声学性能之间，在动力源选择中的声学和耐用性之间，以及在个人飞行的现成硬件的可用性和飞行价值之间。
{"title":"Development of “Aria,” a Compact, Quiet Personal Electric Helicopter","authors":"David A. Coleman, Atanu Halder, Farid Saemi, Carl Runco, H. Denton, Bochan Lee, Vishaal Subramanian, Eric Greenwood, Vinod Lakshminaryan, Moble Benedict","doi":"10.4050/jahs.68.042011","DOIUrl":"https://doi.org/10.4050/jahs.68.042011","url":null,"abstract":"This paper describes the development and flight testing of a personal air vehicle by team Harmony for the GoFly Prize competition. The competition emphasized an open‐air flight experience, and the final aircraft was scored by size (≤8.5 ft), noise (<87 dBA), and speed (>30 kt). The highest scoring team would win a $1 million grand prize. Team Harmony endeavored to develop an aircraft which would maximize the competition score. A counterrotating coaxial electric helicopter configuration was chosen to maximize rotor area and reduce disk loading for efficiency and acoustic benefits. The rotors were designed through a parametric study using an in‐house aerodynamics code, an in‐house acoustics code, and CREATE™‐AV Helios CFD. A quiet electric power train and custom 11 kWh battery were developed. Flight control was implemented with dual, independent, electronically coupled swashplates. A one‐third‐scale prototype aircraft was first developed and flight‐tested. Then a full‐scale, 520 lb (235.4 kg) prototype with an 8.45 ft (2.58 m) rotor diameter was developed and flight‐tested. During hovering, the measured sound pressure levels at 50 ft (15.24 m) were 73 dBA. The full‐scale vehicle crashed before its speed capabilities could be tested, but the aircraft would have likely scored well, achieving the goal of the project. This work encompassed several design trade‐offs, especially between aerodynamic and acoustic performance in the rotor and blade design, and between acoustics and endurance in the power source selection, and in availability and flight‐worthiness of off‐the‐shelf hardware for personal flight.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70220112","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Due to the importance of understanding the aeroacoustics of rotorcraft with continually changing noise sources, this paper presents a new technique for source separation from ground-based acoustic measurements. The source separation process is based on combining a time-domain de-Dopplerization method with the Vold–Kalman order tracking filter approach. This process can extract rotor harmonic noise even when the sources are continuously changing with time, including impulsive events such as blade–vortex interaction noise. The advantage of this approach over traditional methods such as harmonic averaging is that the phase and amplitude relationship of acoustic signals is preserved throughout the extraction process. The approach is applied to the measured acoustic data from a Bell 430 helicopter. The measured data were separated into main rotor harmonic, tail rotor harmonic, and broadband residual components. For steady-state conditions, the extracted components could be depropagated to form acoustic hemispheres showing the directivity of the separated main and tail rotor components. The source separation process was also applied to a maneuvering flight condition. Each component has different pulse shapes and directivity trends, consistent with aeroacoustic theory.
{"title":"Helicopter Noise Source Separation Using an Order Tracking Filter","authors":"Joel Sundar Rachaprolu, Eric Greenwood","doi":"10.4050/jahs.69.012006","DOIUrl":"https://doi.org/10.4050/jahs.69.012006","url":null,"abstract":"Due to the importance of understanding the aeroacoustics of rotorcraft with continually changing noise sources, this paper presents a new technique for source separation from ground-based acoustic measurements. The source separation process is based on combining a time-domain de-Dopplerization method with the Vold–Kalman order tracking filter approach. This process can extract rotor harmonic noise even when the sources are continuously changing with time, including impulsive events such as blade–vortex interaction noise. The advantage of this approach over traditional methods such as harmonic averaging is that the phase and amplitude relationship of acoustic signals is preserved throughout the extraction process. The approach is applied to the measured acoustic data from a Bell 430 helicopter. The measured data were separated into main rotor harmonic, tail rotor harmonic, and broadband residual components. For steady-state conditions, the extracted components could be depropagated to form acoustic hemispheres showing the directivity of the separated main and tail rotor components. The source separation process was also applied to a maneuvering flight condition. Each component has different pulse shapes and directivity trends, consistent with aeroacoustic theory.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135440886","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
In this study, for small unmanned aerial vehicle aeroacoustic predictions, an in-house acoustic code was extended for broadband noise using the empirical Brooks–Pope–Marcolini (BPM) method. The acoustic code was coupled with a comprehensive analysis tool (FLIGHTLAB) or Reynolds-averaged Navier–Stokes (RANS) method to predict the noise from a DJI 9443 CF rotor. The effects of two input parameters of the BPM method on the broadband noise were investigated, including the effective angle of attack and boundary layer parameters. The empirical formula for the boundary layer parameters was replaced with predictions using two-dimensional RANS, allowing arbitrary airfoils other than NACA0012. The boundary layer parameters were also predicted using three-dimensional RANS to capture a three-dimensional flow effect, which can be dominant in modern propellers operating at high rotational speeds. As a result, the three-dimensional effect on the boundary layer was confirmed to conflict with the BPM method, which was developed based on a two-dimensional chordwise flow database. Finally, mid- and high-frequency noise spectra were predicted from the three-dimensional hybrid RANS/LES simulation to be combined with the BPM results, thus improving the noise spectra predictions at midrange frequencies.
{"title":"Empirical Rotor Broadband Noise Prediction Using CFD Boundary Layer Parameter Extraction","authors":"Y. Jung, J. Baeder, Chengjian He","doi":"10.4050/jahs.68.032010","DOIUrl":"https://doi.org/10.4050/jahs.68.032010","url":null,"abstract":"In this study, for small unmanned aerial vehicle aeroacoustic predictions, an in-house acoustic code was extended for broadband noise using the empirical Brooks–Pope–Marcolini (BPM) method. The acoustic code was coupled with a comprehensive analysis tool (FLIGHTLAB) or Reynolds-averaged Navier–Stokes (RANS) method to predict the noise from a DJI 9443 CF rotor. The effects of two input parameters of the BPM method on the broadband noise were investigated, including the effective angle of attack and boundary layer parameters. The empirical formula for the boundary layer parameters was replaced with predictions using two-dimensional RANS, allowing arbitrary airfoils other than NACA0012. The boundary layer parameters were also predicted using three-dimensional RANS to capture a three-dimensional flow effect, which can be dominant in modern propellers operating at high rotational speeds. As a result, the three-dimensional effect on the boundary layer was confirmed to conflict with the BPM method, which was developed based on a two-dimensional chordwise flow database. Finally, mid- and high-frequency noise spectra were predicted from the three-dimensional hybrid RANS/LES simulation to be combined with the BPM results, thus improving the noise spectra predictions at midrange frequencies.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70219267","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Ravi T. Lumba, Cheng Chi, A. Datta, Witold J. F. Koning, Natalia Perez Perez, H. Cummings
The structural design of rotor blades with ultra-thin, unconventional airfoils is conducted in support of the NASA Rotor Optimization for the Advancement of Mars eXploration (ROAMX) project. The outer mold line was provided by NASA, and the internal structural design was developed at the University of Maryland using a CAD-based three-dimensional (3D) aeromechanical analysis. The main objectives of this paper are to document the unique aeroelastic behavior encountered due to the low Reynolds number (down to 15K) and high subsonic Mach number (up to 0.95). Four different blade designs are considered, with the pitch axis varied from quarter-chord to midchord to determine the effect of center of gravity (C.G.) offset on natural frequencies, blade deformations, root loads, and 3D stresses. Torsional stability is emphasized for each of the designs - especially important due to the low Lock number on Mars. The designs are first studied in vacuum, and significant reductions in root loads and 3D stresses are achieved by moving the pitch axis closer to midchord to reduce the C.G. offset. Next, the design with the pitch axis at 40% chord is selected for a lifting-line aeromechanical analysis. The blade control load, airloads, deformations, and 3D stresses are studied for steady hover. Dynamic control load and dynamic 3D stresses are studied for unsteady hover. Interesting elastic twist is observed due to the trapeze effect and propeller moment, in turn affecting the spanwise distribution of aerodynamic loads. The dynamic control load is found to increase significantly due to inertial coupling from the C.G. offset. The dynamic stresses also increase but still have factors of safety greater than two for both tensile and compressive stress.
{"title":"Structural Design and Aeromechanical Analysis of Unconventional Blades for Future Mars Rotorcraft","authors":"Ravi T. Lumba, Cheng Chi, A. Datta, Witold J. F. Koning, Natalia Perez Perez, H. Cummings","doi":"10.4050/jahs.68.042003","DOIUrl":"https://doi.org/10.4050/jahs.68.042003","url":null,"abstract":"The structural design of rotor blades with ultra-thin, unconventional airfoils is conducted in support of the NASA Rotor Optimization for the Advancement of Mars eXploration (ROAMX) project. The outer mold line was provided by NASA, and the internal structural design was developed at the University of Maryland using a CAD-based three-dimensional (3D) aeromechanical analysis. The main objectives of this paper are to document the unique aeroelastic behavior encountered due to the low Reynolds number (down to 15K) and high subsonic Mach number (up to 0.95). Four different blade designs are considered, with the pitch axis varied from quarter-chord to midchord to determine the effect of center of gravity (C.G.) offset on natural frequencies, blade deformations, root loads, and 3D stresses. Torsional stability is emphasized for each of the designs - especially important due to the low Lock number on Mars. The designs are first studied in vacuum, and significant reductions in root loads and 3D stresses are achieved by moving the pitch axis closer to midchord to reduce the C.G. offset. Next, the design with the pitch axis at 40% chord is selected for a lifting-line aeromechanical analysis. The blade control load, airloads, deformations, and 3D stresses are studied for steady hover. Dynamic control load and dynamic 3D stresses are studied for unsteady hover. Interesting elastic twist is observed due to the trapeze effect and propeller moment, in turn affecting the spanwise distribution of aerodynamic loads. The dynamic control load is found to increase significantly due to inertial coupling from the C.G. offset. The dynamic stresses also increase but still have factors of safety greater than two for both tensile and compressive stress.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70219609","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Stefan Platzer, M. Hajek, P. Mortimer, J. Sirohi, Juergen Rauleder
The flow fields of a 2-m diameter two-bladed single rotor, a 2×2-bladed coaxial corotating (stacked) rotor, and a 2×2-bladed coaxial counterrotating (CCR) rotor in hover were measured using particle image velocimetry and computed using a finite-volume unsteady Reynolds-averaged Navier–Stokes (URANS) CFD model. Phase-resolved measurements were performed on the stacked rotor at nine azimuthal locations, and time-resolved measurements were performed on the CCR rotor at 64/rev with at least 500 flow realizations per azimuth for each operating condition. The goal of this study was to compare the flow features of these rotor configurations and explore the interactions between the rotors. Overall, there was good correlation between the measurements and simulations. In particular, the effect of index angle on the upper and lower rotor thrust sharing for the stacked rotor was predicted well by the simulation. The slipstream boundary for the stacked rotor was found to vary with the index angle. The slipstream boundary and vortex trajectories for the CCR rotor were found to vary with azimuthal location, indicating the effect of blade passage on the wake geometry. Simulations indicated a stronger dependence of the tip vortex trajectory on the index angle and thrust for the stacked rotor compared to the CCR rotor. The radial thrust distribution along the upper blades was found to depend on the index angle for the stacked rotor and showed small variation due to blade passage for the CCR rotor. A larger azimuthal dependence was seen for the radial thrust distribution on the lower rotor blades, primarily due to the proximity of the upper rotor tip vortices. The lower rotor radial thrust distribution was biased towards the blade tip, outside the upper rotor slipstream.
{"title":"Investigation of the Flow Fields of Coaxial Co-Rotating and Counter-Rotating Rotors in Hover Using Measurements and Simulations","authors":"Stefan Platzer, M. Hajek, P. Mortimer, J. Sirohi, Juergen Rauleder","doi":"10.4050/jahs.68.042008","DOIUrl":"https://doi.org/10.4050/jahs.68.042008","url":null,"abstract":"The flow fields of a 2-m diameter two-bladed single rotor, a 2×2-bladed coaxial corotating (stacked) rotor, and a 2×2-bladed coaxial counterrotating (CCR) rotor in hover were measured using particle image velocimetry and computed using a finite-volume unsteady Reynolds-averaged Navier–Stokes (URANS) CFD model. Phase-resolved measurements were performed on the stacked rotor at nine azimuthal locations, and time-resolved measurements were performed on the CCR rotor at 64/rev with at least 500 flow realizations per azimuth for each operating condition. The goal of this study was to compare the flow features of these rotor configurations and explore the interactions between the rotors. Overall, there was good correlation between the measurements and simulations. In particular, the effect of index angle on the upper and lower rotor thrust sharing for the stacked rotor was predicted well by the simulation. The slipstream boundary for the stacked rotor was found to vary with the index angle. The slipstream boundary and vortex trajectories for the CCR rotor were found to vary with azimuthal location, indicating the effect of blade passage on the wake geometry. Simulations indicated a stronger dependence of the tip vortex trajectory on the index angle and thrust for the stacked rotor compared to the CCR rotor. The radial thrust distribution along the upper blades was found to depend on the index angle for the stacked rotor and showed small variation due to blade passage for the CCR rotor. A larger azimuthal dependence was seen for the radial thrust distribution on the lower rotor blades, primarily due to the proximity of the upper rotor tip vortices. The lower rotor radial thrust distribution was biased towards the blade tip, outside the upper rotor slipstream.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70219397","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
University of Maryland's Maryland Tiltrotor Rig was tested in the Naval Surface Warfare Center Carderock Division 8-×10-ft subsonic wind tunnel. Flutter frequency and damping data were collected for wing beam and chord modes up to 100 kt. Eight configurations were tested. Baseline data are gimbal-free, freewheeling mode, wing fairings on with straight and swept-tip blades. Gimbal-locked, powered mode, and wing fairings off data were also collected, all with straight and swept-tip blades. The sweep angle is 20°, starting at 80%R. Details of the mathematical model are reported. Predictions were carried out for each configuration with the University of Maryland's new aeromechanics solver UMARC-II. Wing beam damping showed good agreement with the test data. Wing chord damping was underpredicted with a maximum of 0.9% difference. The trends for this mode for the gimbal-locked, straight blades configurations (freewheeling and powered) were not captured by the analysis. Swept-tip blades did not show a definitive increase in wing beam or chord damping for the gimbal-free configuration. However, wing chord damping increased (about 0.4% at 60 kt) due to swept-tip blades for the gimbal-locked, freewheeling configuration. Locking the gimbal increased the wing chord damping by 0.5%, which was picked up by the analysis. Powered mode also increased the wing chord damping by 0.5% compared to freewheeling mode, but the analysis did not predict this behavior. Wing beam damping test data showed an increase at higher speeds due to wing aerodynamics, although not as clearly as predictions due to scatter.
{"title":"Prediction and Validation of Whirl Flutter Data of the Maryland Tiltrotor Rig","authors":"Seyhan Gul, Anubhav Datta","doi":"10.4050/jahs.69.012011","DOIUrl":"https://doi.org/10.4050/jahs.69.012011","url":null,"abstract":"University of Maryland's Maryland Tiltrotor Rig was tested in the Naval Surface Warfare Center Carderock Division 8-×10-ft subsonic wind tunnel. Flutter frequency and damping data were collected for wing beam and chord modes up to 100 kt. Eight configurations were tested. Baseline data are gimbal-free, freewheeling mode, wing fairings on with straight and swept-tip blades. Gimbal-locked, powered mode, and wing fairings off data were also collected, all with straight and swept-tip blades. The sweep angle is 20°, starting at 80%R. Details of the mathematical model are reported. Predictions were carried out for each configuration with the University of Maryland's new aeromechanics solver UMARC-II. Wing beam damping showed good agreement with the test data. Wing chord damping was underpredicted with a maximum of 0.9% difference. The trends for this mode for the gimbal-locked, straight blades configurations (freewheeling and powered) were not captured by the analysis. Swept-tip blades did not show a definitive increase in wing beam or chord damping for the gimbal-free configuration. However, wing chord damping increased (about 0.4% at 60 kt) due to swept-tip blades for the gimbal-locked, freewheeling configuration. Locking the gimbal increased the wing chord damping by 0.5%, which was picked up by the analysis. Powered mode also increased the wing chord damping by 0.5% compared to freewheeling mode, but the analysis did not predict this behavior. Wing beam damping test data showed an increase at higher speeds due to wing aerodynamics, although not as clearly as predictions due to scatter.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135505053","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper covers the design, fabrication, testing, and modeling of a family of Froude-scale tiltrotor blades. They are designed with the objective of gaining a fundamental understanding of the impact of a swept tip on tiltrotor whirl flutter. The goal of this paper is to describe the development of the blades needed for this purpose. The rotor is three bladed with a diameter of 4.75 ft. The blades have a VR-7 profile, chord of 3.15 inches, and linear twist of –37° per span. The swept-tip blades have a sweep of 20° starting at 80% R . The blade properties are loosely based on the XV-15 design. A CATIA and Cubit-based high-fidelity three-dimensional (3D) finite element model is developed. It accurately represents the fabricated blade and is analyzed with X3D. Experiments in a vacuum chamber were carried out to demonstrate the structural integrity of the blades. Measured frequencies and strains were validated with X3D predictions proving the fidelity of the 3D model. Thus, even though the wind tunnel facilities were closed due to COVID-19, hover and forward flight calculations for the blade stress could be performed using the high-fidelity 3D structural model. The results prove the blades have sufficient structural integrity and stress margins to allow for wind tunnel testing.
{"title":"Fabrication, Testing, and 3D Comprehensive Analysis of Swept-Tip Tiltrotor Blades","authors":"James Sutherland, Anubhav Datta","doi":"10.4050/jahs.68.012002","DOIUrl":"https://doi.org/10.4050/jahs.68.012002","url":null,"abstract":"This paper covers the design, fabrication, testing, and modeling of a family of Froude-scale tiltrotor blades. They are designed with the objective of gaining a fundamental understanding of the impact of a swept tip on tiltrotor whirl flutter. The goal of this paper is to describe the development of the blades needed for this purpose. The rotor is three bladed with a diameter of 4.75 ft. The blades have a VR-7 profile, chord of 3.15 inches, and linear twist of –37° per span. The swept-tip blades have a sweep of 20° starting at 80% R . The blade properties are loosely based on the XV-15 design. A CATIA and Cubit-based high-fidelity three-dimensional (3D) finite element model is developed. It accurately represents the fabricated blade and is analyzed with X3D. Experiments in a vacuum chamber were carried out to demonstrate the structural integrity of the blades. Measured frequencies and strains were validated with X3D predictions proving the fidelity of the 3D model. Thus, even though the wind tunnel facilities were closed due to COVID-19, hover and forward flight calculations for the blade stress could be performed using the high-fidelity 3D structural model. The results prove the blades have sufficient structural integrity and stress margins to allow for wind tunnel testing.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135077878","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
A hingeless hub tiltrotor with swept-tip blades was examined comprehensively with a new rotorcraft aeromechanics solver developed at the University of Maryland. The solver was verified with hypothetical U.S. Army results and validated with Boeing Model 222 test data from 1972. A 20° sweep back from 80%R increased instability speed to 405 kt, an improvement of more than 75 kt. The key mechanism is the aerodynamic center shift. The trade-off is the increase in control system and blade loads. Fundamental understanding of physics is provided. Proprotor air resonance emerged as the critical phenomenon, not whirl flutter. Predictions in powered mode are necessary. At least the first rotor flap, lag, and torsion modes need to be included. Rotor aerodynamics should use airfoil tables; wing aerodynamics is not important for air resonance. Analysis shows high-speed flight is achievable with 13.5% thick wings but systematic wind tunnel tests with modern equipment are necessary for further validation.
{"title":"Aeroelastic Loads and Stability of Swept-Tip Hingeless Tiltrotors toward High-Speed Instability-Free Cruise","authors":"Seyhan Gul, A. Datta","doi":"10.4050/jahs.68.012001","DOIUrl":"https://doi.org/10.4050/jahs.68.012001","url":null,"abstract":"A hingeless hub tiltrotor with swept-tip blades was examined comprehensively with a new rotorcraft aeromechanics solver developed at the University of Maryland. The solver was verified with hypothetical U.S. Army results and validated with Boeing Model 222 test data from 1972. A 20° sweep back from 80%R increased instability speed to 405 kt, an improvement of more than 75 kt. The key mechanism is the aerodynamic center shift. The trade-off is the increase in control system and blade loads. Fundamental understanding of physics is provided. Proprotor air resonance emerged as the critical phenomenon, not whirl flutter. Predictions in powered mode are necessary. At least the first rotor flap, lag, and torsion modes need to be included. Rotor aerodynamics should use airfoil tables; wing aerodynamics is not important for air resonance. Analysis shows high-speed flight is achievable with 13.5% thick wings but systematic wind tunnel tests with modern equipment are necessary for further validation.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":null,"pages":null},"PeriodicalIF":1.5,"publicationDate":"2022-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70218389","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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