M. May, N. Schneider, B. Schaufelberger, Markus Jung, Jonas Pfaff, Anja Altes, T. Haase, S. Schopferer, M. Imbert, P. Matura, Martin Blacha
Drones operating in an urban environment pose a potential collision threat to rotorcraft. In this paper, the battery pack of the DJI MAVIC 2 ZOOM is analyzed since the battery is considered to be the greatest threat due to its high weight and stiffness. Following a pyramid-type building block approach, a high-fidelity model simulation was developed for LS-DYNA based on a wide range of experiments, ranging from quasi-static material tests to quasi-static component tests up to high-velocity impact experiments. The high-fidelity model allows the prediction of damage in potential collision scenarios between a high-speed rotorcraft and the battery pack of the drone. For the particular impact configuration analyzed within this paper, the drone battery does not cause catastrophic failure of the windshield of the rotorcraft.
{"title":"Collisions between Drones and Rotorcraft: Modeling of the Crash Response of Battery Packs","authors":"M. May, N. Schneider, B. Schaufelberger, Markus Jung, Jonas Pfaff, Anja Altes, T. Haase, S. Schopferer, M. Imbert, P. Matura, Martin Blacha","doi":"10.4050/jahs.69.012004","DOIUrl":"https://doi.org/10.4050/jahs.69.012004","url":null,"abstract":"Drones operating in an urban environment pose a potential collision threat to rotorcraft. In this paper, the battery pack of the DJI MAVIC 2 ZOOM is analyzed since the battery is considered to be the greatest threat due to its high weight and stiffness. Following a pyramid-type building block approach, a high-fidelity model simulation was developed for LS-DYNA based on a wide range of experiments, ranging from quasi-static material tests to quasi-static component tests up to high-velocity impact experiments. The high-fidelity model allows the prediction of damage in potential collision scenarios between a high-speed rotorcraft and the battery pack of the drone. For the particular impact configuration analyzed within this paper, the drone battery does not cause catastrophic failure of the windshield of the rotorcraft.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"44 1","pages":""},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70219706","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}
Ting-Wei Hsu, Jason J. Choi, Divyang Amin, Claire Tomlin, Shaun C. McWherter, Michael Piedmonte
Innovative electric vertical take-off and landing (eVTOL) aircraft designs and operational concepts, driven by advancements in battery and electric motor technologies, seek to achieve superior safety records with increased system redundancy. Validating safe flight operations within verified flight envelope regions for passenger flights in densely populated urban environments remains a primary challenge. This paper establishes a framework for applying Hamilton–Jacobi reachability analysis to the full six-degree-of-freedom (6-DOF) dynamics of the NASA Tiltwing vehicle, verifying the flight envelope during the flight mode transition between near-hover and cruise flight, which prevents loss of control of the vehicle and ensures recoverability to safe trim conditions. This involves first verifying the nominal flight mode transition path as a series of trim points, defining the safe flight envelope using reachability, and decomposing the system dynamics into longitudinal and lateral subsystems. Our formulation guarantees the computed envelope's robustness against modeling errors and uncertainties, and the usage of state decomposition significantly improves the tractability of the reachability computation. The result is validated through Monte Carlo 6-DOF nonlinear simulation of vehicle dynamics, demonstrating that the vehicle states within the flight envelope can successfully recover to trim states and continue the flight mode transition safely.
{"title":"Towards Flight Envelope Protection for the NASA Tiltwing eVTOL Flight Mode Transition Using Hamilton–Jacobi Reachability","authors":"Ting-Wei Hsu, Jason J. Choi, Divyang Amin, Claire Tomlin, Shaun C. McWherter, Michael Piedmonte","doi":"10.4050/jahs.69.022003","DOIUrl":"https://doi.org/10.4050/jahs.69.022003","url":null,"abstract":"Innovative electric vertical take-off and landing (eVTOL) aircraft designs and operational concepts, driven by advancements in battery and electric motor technologies, seek to achieve superior safety records with increased system redundancy. Validating safe flight operations within verified flight envelope regions for passenger flights in densely populated urban environments remains a primary challenge. This paper establishes a framework for applying Hamilton–Jacobi reachability analysis to the full six-degree-of-freedom (6-DOF) dynamics of the NASA Tiltwing vehicle, verifying the flight envelope during the flight mode transition between near-hover and cruise flight, which prevents loss of control of the vehicle and ensures recoverability to safe trim conditions. This involves first verifying the nominal flight mode transition path as a series of trim points, defining the safe flight envelope using reachability, and decomposing the system dynamics into longitudinal and lateral subsystems. Our formulation guarantees the computed envelope's robustness against modeling errors and uncertainties, and the usage of state decomposition significantly improves the tractability of the reachability computation. The result is validated through Monte Carlo 6-DOF nonlinear simulation of vehicle dynamics, demonstrating that the vehicle states within the flight envelope can successfully recover to trim states and continue the flight mode transition safely.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"71 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135214617","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 describes the implementation and the evaluation of newly designed helicopter ship deck landing control modes in a piloted simulation study. The ship deck landing modes are embedded in a model-following controller architecture. The employed control design is a complete model-following control system, which imposes the desired command model dynamics on the controlled helicopter. Different command types combined with various hold functions are implemented to make the task easier for the pilots. Three basic command modes and three advanced command modes, one without ship communication and two with ship communication, are implemented. A piloted simulation study was performed in a simulator to evaluate and compare the implemented control modes within a complete maritime scenario design. The evaluation of control modes is based on the success of helicopter ship deck landings which is assessed by a quantitative as well as a qualitative assessment. Simulation results demonstrate that the advanced command modes improved the task performance as well as reduced the pilot workload extensively in comparison to the basic command modes.
{"title":"Evaluation of Helicopter Ship Deck Landing Control Laws in Piloted Simulations","authors":"Arti Kalra, Alexander Štrbac, Malte-Jörn Maibach","doi":"10.4050/jahs.69.012002","DOIUrl":"https://doi.org/10.4050/jahs.69.012002","url":null,"abstract":"This paper describes the implementation and the evaluation of newly designed helicopter ship deck landing control modes in a piloted simulation study. The ship deck landing modes are embedded in a model-following controller architecture. The employed control design is a complete model-following control system, which imposes the desired command model dynamics on the controlled helicopter. Different command types combined with various hold functions are implemented to make the task easier for the pilots. Three basic command modes and three advanced command modes, one without ship communication and two with ship communication, are implemented. A piloted simulation study was performed in a simulator to evaluate and compare the implemented control modes within a complete maritime scenario design. The evaluation of control modes is based on the success of helicopter ship deck landings which is assessed by a quantitative as well as a qualitative assessment. Simulation results demonstrate that the advanced command modes improved the task performance as well as reduced the pilot workload extensively in comparison to the basic command modes.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136306063","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 investigates the interactional aerodynamics of hovering side-by-side rotors in ground effect. The 5.5-ft diameter, three-bladed fixed-pitched rotors are simulated using computational fluid dynamics at a targeted 5 lb/ft2 disk loading. Simulations are performed using the commercial Navier–Stokes solver, AcuSolve®, with a delayed detached eddy simulation model. Side-by-side rotors are simulated at two heights above the ground (H/D = 0.5 and H/D = 1), and with two hub–hub separation distances (3R and 2.5R). The performance of side-by-side rotors in ground effect (IGE) is compared to isolated rotors out of ground effect. Between the side-by-side rotors IGE, a highly turbulent mixing region is identified where the wakes of each rotor collide. The flow fountains upwards, as well as exits outwards (along a direction normal to a plane connecting the two rotor hubs) with fountaining between the rotors reaching up to 0.75D above the ground. As blades at H/D = 0.5 traverse the highly turbulent flow, strong vibratory loading is induced and a thrust loss is observed over the outboard sections between the rotors that is large enough to negate any nominal ground effect benefits inboard. Side-by-side rotors at H/D = 0.5 with 2.5R hub–hub spacing produce peak-to-peak thrust oscillations up to 16% of the steady thrust. Rotors placed higher, at H/D = 1 are positioned above the turbulent mixing flow and produce significantly lower vibratory loads. The spacing between rotors at H/D = 0.5 and 3R hub–hub separation allows strong vortical structures to develop between the rotors which move from side to side over multiple revolutions. When the vorticity moves closer to one of the rotors, it produces a greater lift deficit over the outboard region and a stronger vibratory loading. For rotors closer together, at H/D = 0.5 and 2.5R separation, the vortical structures between rotors are constrained to a more concentrated area and show less side-to-side drift.
{"title":"A Computational Examination of Side-by-Side Rotors in Ground Effect","authors":"R. Healy, J. McCauley, F. Gandhi, O. Sahni","doi":"10.4050/jahs.68.032007","DOIUrl":"https://doi.org/10.4050/jahs.68.032007","url":null,"abstract":"This study investigates the interactional aerodynamics of hovering side-by-side rotors in ground effect. The 5.5-ft diameter, three-bladed fixed-pitched rotors are simulated using computational fluid dynamics at a targeted 5 lb/ft2 disk loading. Simulations are performed using the commercial Navier–Stokes solver, AcuSolve®, with a delayed detached eddy simulation model. Side-by-side rotors are simulated at two heights above the ground (H/D = 0.5 and H/D = 1), and with two hub–hub separation distances (3R and 2.5R). The performance of side-by-side rotors in ground effect (IGE) is compared to isolated rotors out of ground effect. Between the side-by-side rotors IGE, a highly turbulent mixing region is identified where the wakes of each rotor collide. The flow fountains upwards, as well as exits outwards (along a direction normal to a plane connecting the two rotor hubs) with fountaining between the rotors reaching up to 0.75D above the ground. As blades at H/D = 0.5 traverse the highly turbulent flow, strong vibratory loading is induced and a thrust loss is observed over the outboard sections between the rotors that is large enough to negate any nominal ground effect benefits inboard. Side-by-side rotors at H/D = 0.5 with 2.5R hub–hub spacing produce peak-to-peak thrust oscillations up to 16% of the steady thrust. Rotors placed higher, at H/D = 1 are positioned above the turbulent mixing flow and produce significantly lower vibratory loads. The spacing between rotors at H/D = 0.5 and 3R hub–hub separation allows strong vortical structures to develop between the rotors which move from side to side over multiple revolutions. When the vorticity moves closer to one of the rotors, it produces a greater lift deficit over the outboard region and a stronger vibratory loading. For rotors closer together, at H/D = 0.5 and 2.5R separation, the vortical structures between rotors are constrained to a more concentrated area and show less side-to-side drift.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"1 1","pages":""},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70218696","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 article describes the implementation and linearization of free-vortex wake models in state-variable form as applied to rotary- and flapping-wing vehicles. More specifically, the wake models are implemented and tested for a UH-60 rotor in forward flight and for a hovering insect representative of a hawk moth. A periodic solution to each wake model is found by time marching the coupled rotor/wing and vortex wake dynamics. Next, linearized harmonic decomposition models are obtained and validated against nonlinear simulations. Order reduction methods are explored to guide the development of linearized wake models that provide increased runtime performance compared to the nonlinear and linearized harmonic decomposition wake models while guaranteeing satisfactory prediction of the periodic response of the wake. This constitutes a first attempt to extend free-vortex wake methods in state-variable form, originally developed for rotary-wing applications, to flapping-wing flight.
{"title":"Implementation and Linearization of State-Space Free-Vortex Wake Models for Rotary- and Flapping-Wing Vehicles","authors":"Umberto Saetti, J. Horn","doi":"10.4050/jahs.68.042004","DOIUrl":"https://doi.org/10.4050/jahs.68.042004","url":null,"abstract":"This article describes the implementation and linearization of free-vortex wake models in state-variable form as applied to rotary- and flapping-wing vehicles. More specifically, the wake models are implemented and tested for a UH-60 rotor in forward flight and for a hovering insect representative of a hawk moth. A periodic solution to each wake model is found by time marching the coupled rotor/wing and vortex wake dynamics. Next, linearized harmonic decomposition models are obtained and validated against nonlinear simulations. Order reduction methods are explored to guide the development of linearized wake models that provide increased runtime performance compared to the nonlinear and linearized harmonic decomposition wake models while guaranteeing satisfactory prediction of the periodic response of the wake. This constitutes a first attempt to extend free-vortex wake methods in state-variable form, originally developed for rotary-wing applications, to flapping-wing flight.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"1 1","pages":""},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70219252","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}
Recent advances in urban air mobility have driven the development of many new vertical take-off and landing (VTOL) concepts. These vehicles often feature original designs departing from the conventional helicopter configuration. Due to their novelty, the characteristics of the supervortices forming in the wake of such aircraft are unknown. However, these vortices may endanger any other vehicle evolving in their close proximity, owing to potentially large induced velocities. Therefore, improved knowledge about the wakes of VTOL vehicles is needed to guarantee safe urban air mobility operations. In this work, we study the wake of three VTOL aircraft in cruise by means of large eddy simulation. We present a two-stage numerical procedure that enables the simulation of long wake ages at a limited computational cost. Our simulations reveal that the wakes of rotary vehicles (quadcopter and side-by-side helicopter) feature larger wake vortex cores than an isolated wing. Their decay is also accelerated due to self-induced turbulence generated during the wake roll-up. A tilt-wing wake, on the other hand, is moderately turbulent and has smaller vortex cores than the wing. Finally, we introduce an empirical model of the vortex circulation distribution that enables fast prediction of wake-induced velocities, within a 2% error of the simulation results on average.
{"title":"Large Eddy Simulation for Empirical Modeling of the Wake of Three Urban Air Mobility Vehicles","authors":"Denis-Gabriel Caprace, A. Ning","doi":"10.4050/jahs.68.042002","DOIUrl":"https://doi.org/10.4050/jahs.68.042002","url":null,"abstract":"Recent advances in urban air mobility have driven the development of many new vertical take-off and landing (VTOL) concepts. These vehicles often feature original designs departing from the conventional helicopter configuration. Due to their novelty, the characteristics of the supervortices forming in the wake of such aircraft are unknown. However, these vortices may endanger any other vehicle evolving in their close proximity, owing to potentially large induced velocities. Therefore, improved knowledge about the wakes of VTOL vehicles is needed to guarantee safe urban air mobility operations. In this work, we study the wake of three VTOL aircraft in cruise by means of large eddy simulation. We present a two-stage numerical procedure that enables the simulation of long wake ages at a limited computational cost. Our simulations reveal that the wakes of rotary vehicles (quadcopter and side-by-side helicopter) feature larger wake vortex cores than an isolated wing. Their decay is also accelerated due to self-induced turbulence generated during the wake roll-up. A tilt-wing wake, on the other hand, is moderately turbulent and has smaller vortex cores than the wing. Finally, we introduce an empirical model of the vortex circulation distribution that enables fast prediction of wake-induced velocities, within a 2% error of the simulation results on average.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"1 1","pages":""},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70219532","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}
The effects of ambient atmospheric conditions, air temperature, and density on rotor harmonic noise radiation are characterized using theoretical models and experimental measurements of helicopter noise collected at three different test sites at elevations ranging from sea level to 7000 ft above sea level. Significant changes in the thickness, loading, and blade–vortex interaction noise levels and radiation directions are observed across the different test sites for an AS350 helicopter flying at the same indicated airspeed and gross weight. However, the radiated noise is shown to scale with ambient pressure when the flight condition of the helicopter is defined in nondimensional terms. Although the effective tip Mach number is identified as the primary governing parameter for thickness noise, the nondimensional weight coefficient also impacts lower harmonic loading noise levels, which contribute strongly to low-frequency harmonic noise radiation both in and out of the plane of the horizon. Strategies for maintaining the same nondimensional rotor operating condition under different ambient conditions are developed using an analytical model of single main rotor helicopter trim and confirmed using a CAMRAD II model of the AS350 helicopter. The ability of the Fundamental Rotorcraft Acoustics Modeling from Experiments (FRAME) technique to generalize noise measurements made under one set of ambient conditions to make accurate noise predictions under other ambient conditions is also validated.
{"title":"Effects of Ambient Conditions on Helicopter Harmonic Noise Radiation: Theory and Experiment","authors":"Eric Greenwood, Ben W. Sim, D. D. Boyd, Jr.","doi":"10.4050/jahs.69.012005","DOIUrl":"https://doi.org/10.4050/jahs.69.012005","url":null,"abstract":"The effects of ambient atmospheric conditions, air temperature, and density on rotor harmonic noise radiation are characterized using theoretical models and experimental measurements of helicopter noise collected at three different test sites at elevations ranging from sea level to 7000 ft above sea level. Significant changes in the thickness, loading, and blade–vortex interaction noise levels and radiation directions are observed across the different test sites for an AS350 helicopter flying at the same indicated airspeed and gross weight. However, the radiated noise is shown to scale with ambient pressure when the flight condition of the helicopter is defined in nondimensional terms. Although the effective tip Mach number is identified as the primary governing parameter for thickness noise, the nondimensional weight coefficient also impacts lower harmonic loading noise levels, which contribute strongly to low-frequency harmonic noise radiation both in and out of the plane of the horizon. Strategies for maintaining the same nondimensional rotor operating condition under different ambient conditions are developed using an analytical model of single main rotor helicopter trim and confirmed using a CAMRAD II model of the AS350 helicopter. The ability of the Fundamental Rotorcraft Acoustics Modeling from Experiments (FRAME) technique to generalize noise measurements made under one set of ambient conditions to make accurate noise predictions under other ambient conditions is also validated.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"1 1","pages":""},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70219808","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}
Paolo Francesco Scaramuzzino, Daan M. Pool, Marilena D. Pavel, Olaf Stroosma, Giuseppe Quaranta, Max Mulder
This paper analyzes the effects of the helicopter dynamics on pilots' learning process and transfer of learned skills during autorotation training. A quasi-transfer-of-training experiment was performed with 10 experienced helicopter pilots in the SIMONA moving-base flight simulator at Delft University of Technology. Pilots had to control an in-house flight dynamics model setup to simulate two types of helicopter dynamics: (1) a “hard” dynamics characterized by a low autorotative flare index requiring high pilot control compensation and (2) a “easy” dynamics characterized by a high autorotative flare index with low pilot control compensation required. Two groups of pilots tested these types of dynamics in a different training sequence: hard-easy-hard (HEH group) and easy-hard-easy (EHE group). The main conclusion of this study proved that simulator training for autorotation can best start with pilots training in the most resource demanding condition. A more challenging helicopter's dynamics will require higher pilot agility and more rapid responses to his/her perceptual changes. This will result in pilots developing more robust and adaptable flying skills. Indeed, a clear positive transfer of training effect was observed in the experiment presented in this paper in terms of acquired pilot skills in the HEH group, but not the EHE group. Positive transfer was especially observed in terms of reduced rate of descent at touchdown. The two groups differed in the control strategy applied, with the HEH group having developed a control technique mimicking more closely the one adopted in a real helicopter.
{"title":"Autorotation Transfer of Training: Effects of Helicopter Dynamics","authors":"Paolo Francesco Scaramuzzino, Daan M. Pool, Marilena D. Pavel, Olaf Stroosma, Giuseppe Quaranta, Max Mulder","doi":"10.4050/jahs.69.022007","DOIUrl":"https://doi.org/10.4050/jahs.69.022007","url":null,"abstract":"This paper analyzes the effects of the helicopter dynamics on pilots' learning process and transfer of learned skills during autorotation training. A quasi-transfer-of-training experiment was performed with 10 experienced helicopter pilots in the SIMONA moving-base flight simulator at Delft University of Technology. Pilots had to control an in-house flight dynamics model setup to simulate two types of helicopter dynamics: (1) a “hard” dynamics characterized by a low autorotative flare index requiring high pilot control compensation and (2) a “easy” dynamics characterized by a high autorotative flare index with low pilot control compensation required. Two groups of pilots tested these types of dynamics in a different training sequence: hard-easy-hard (HEH group) and easy-hard-easy (EHE group). The main conclusion of this study proved that simulator training for autorotation can best start with pilots training in the most resource demanding condition. A more challenging helicopter's dynamics will require higher pilot agility and more rapid responses to his/her perceptual changes. This will result in pilots developing more robust and adaptable flying skills. Indeed, a clear positive transfer of training effect was observed in the experiment presented in this paper in terms of acquired pilot skills in the HEH group, but not the EHE group. Positive transfer was especially observed in terms of reduced rate of descent at touchdown. The two groups differed in the control strategy applied, with the HEH group having developed a control technique mimicking more closely the one adopted in a real helicopter.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"22 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135505049","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}
The first whirl flutter test of the Maryland Tiltrotor Rig (MTR) was recently completed in the Naval Surface Warfare Center Carderock Division (NSWCCD) 8- by 10-ft large subsonic wind tunnel. The MTR is a Froude-scaled, 4.75 ft diameter, three-bladed, semispan, floor-mounted, optionally powered, flutter rig. This paper focuses on the swept-tip blades developed for this test. The swept-tip begins at 80%R and sweeps aft 20° to alleviate whirl flutter. Whirl flutter testing was performed in four configurations for both blade geometries. The frequency and damping of the wing beam and chord bending were collected at wind speeds up to 100 kt. This paper directly compares whirl flutter results between the straight and swept-tip blades and demonstrates higher wing chord damping using swept-tip blades in freewheel conditions, even at lower flight speeds.
{"title":"Whirl Flutter Test of Swept-Tip Tiltrotor Blades","authors":"James Sutherland, Frederick Tsai, Anubhav Datta","doi":"10.4050/jahs.69.012010","DOIUrl":"https://doi.org/10.4050/jahs.69.012010","url":null,"abstract":"The first whirl flutter test of the Maryland Tiltrotor Rig (MTR) was recently completed in the Naval Surface Warfare Center Carderock Division (NSWCCD) 8- by 10-ft large subsonic wind tunnel. The MTR is a Froude-scaled, 4.75 ft diameter, three-bladed, semispan, floor-mounted, optionally powered, flutter rig. This paper focuses on the swept-tip blades developed for this test. The swept-tip begins at 80%R and sweeps aft 20° to alleviate whirl flutter. Whirl flutter testing was performed in four configurations for both blade geometries. The frequency and damping of the wing beam and chord bending were collected at wind speeds up to 100 kt. This paper directly compares whirl flutter results between the straight and swept-tip blades and demonstrates higher wing chord damping using swept-tip blades in freewheel conditions, even at lower flight speeds.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"10 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135505038","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}
Elena Shrestha, Derrick W. Yeo, Vikram Hrishikeshavan, I. Chopra
This paper describes the gust rejection study of a twin-cyclocopter micro air vehicle with two cyclorotors and an anti-torque nose rotor. A gust rejection controller relying on velocity feedback and onboard flow sensing was implemented in a closed-loop feedback system. Tethered experiments were conducted with the vehicle mounted on a 6-DOF stand in front of a synthetic gust generation device capable of providing up to 4 m/s step gust input. Planar gusts along the longitudinal and lateral axis of the cyclocopter and crosswinds were systematically studied. Both pitch control (varying nose rotor rpm) and thrust vectoring control strategies were evaluated to counteract perturbation along the longitudinal axis, while the cyclocopter used differential rotational speed of the cyclorotors for lateral gusts. Results showed that thrust vectoring was more effective than pitch control in reducing displacement from gust. Pitch control using the position feedback controller resulted in a maximum gust tolerance of 2.8 m/s with a duration of 3 s. while thrust vector control using the combined flow feedback and position feedback controller was able to withstand 4 m/s step gust input with 1 s duration with only 0.01 m displacement. In addition, flow feedback was more effective than position feedback in minimizing position error. The cyclocopter was also able to mitigate step gusts of 2.8 m/s in magnitude and 3 s duration with crosswind components at 30 deg from the longitudinal axis. The difference in performance can be attributed to the cyclocopter's unique thrust vectoring capability.
{"title":"Improved Gust Tolerance of a Cyclocopter UAV Using Onboard Flow Sensing","authors":"Elena Shrestha, Derrick W. Yeo, Vikram Hrishikeshavan, I. Chopra","doi":"10.4050/jahs.68.032011","DOIUrl":"https://doi.org/10.4050/jahs.68.032011","url":null,"abstract":"This paper describes the gust rejection study of a twin-cyclocopter micro air vehicle with two cyclorotors and an anti-torque nose rotor. A gust rejection controller relying on velocity feedback and onboard flow sensing was implemented in a closed-loop feedback system. Tethered experiments were conducted with the vehicle mounted on a 6-DOF stand in front of a synthetic gust generation device capable of providing up to 4 m/s step gust input. Planar gusts along the longitudinal and lateral axis of the cyclocopter and crosswinds were systematically studied. Both pitch control (varying nose rotor rpm) and thrust vectoring control strategies were evaluated to counteract perturbation along the longitudinal axis, while the cyclocopter used differential rotational speed of the cyclorotors for lateral gusts. Results showed that thrust vectoring was more effective than pitch control in reducing displacement from gust. Pitch control using the position feedback controller resulted in a maximum gust tolerance of 2.8 m/s with a duration of 3 s. while thrust vector control using the combined flow feedback and position feedback controller was able to withstand 4 m/s step gust input with 1 s duration with only 0.01 m displacement. In addition, flow feedback was more effective than position feedback in minimizing position error. The cyclocopter was also able to mitigate step gusts of 2.8 m/s in magnitude and 3 s duration with crosswind components at 30 deg from the longitudinal axis. The difference in performance can be attributed to the cyclocopter's unique thrust vectoring capability.","PeriodicalId":50017,"journal":{"name":"Journal of the American Helicopter Society","volume":"1 1","pages":""},"PeriodicalIF":1.5,"publicationDate":"2023-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70219367","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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