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This paper focused on safety considerations in Urban Air Mobility (UAM) through a cross-industry examination of commercial aviation, rotorcraft, and unmanned aerial systems (UAS). Although UAM promises transformative benefits, there are safety concerns remaining. Based on the Federal Aviation Administration (FAA)’s Concept of Operations (ConOps), the literature review explained the fundamental concepts of UAM. In commercial aviation, regulatory framework, pilot training and certification, vehicle design and maintenance, and emergency response planning are emphasized. For rotorcraft, safety requirements for vertical flight, collision avoidance systems, heliport standards, and weather adaptability are crucial. Leveraging UAS advancements, the study suggested autonomous systems, sense-and-avoid technology, and remote piloting for enhanced safety in the UAM sector.

The aim of this paper is to provide a survey on the validation metrics and criteria employed for validation of rotorcraft modeling accuracy, extending the present approaches to the uncertainty quantification (UQ) domain of validation. In this sense, the paper will review the current state-of-the-art metrics belonging to 1) the root mean square error method which are expressed as cost functions in frequency and time domain and 2) their enhancement towards the chi-square statistics consideration. The paper will extend afterwards the problem of validation towards uncertainty quantification (UQ) area with its distinction between aleatory and epistemic uncertainties. In UQ the emphasis is not on validation metrics but on the methods to validate the model, the choice of a metric depending on the problem under investigation. A simple example will be given for rotorcraft showing that the problem of flight controller modeling can become nondeterministic using different formulations for the equations of motion. A key message of the paper is that using UQ in the future can enhance our understanding of the simulation model fidelity.

Purpose: The purpose of this study was to examine the noise profiles of helicopters and eVTOLs at heliports and vertiports in order to support site planning and the development of local land-use compatibility guidelines. Design/methodology/approach: This study used existing helicopter and eVTOL sound profiles to develop noise distribution contours for heliports and vertiports. This study collected noise data from available studies on helicopters and eVTOLs. This data was then entered into dBmap noise mapping tool to develop noise maps for helicopters and eVTOLs. Three air transit facility configurations for each aircraft category were analyzed: open (unobstructed), urban, and rooftop. Two additional vertiport configurations were created by overlaying building and noise data with Google Earth imagery. Noise data was entered into dBmap for processing. Noise contours were mapped for three configurations for both aircraft categories. Findings: Key observations from the noise distribution analysis showed that open (unobstructed) facility configurations have even noise distributions for both eVTOLs and helicopters. Noise maps for urban locations showed areas of attenuation and asymmetrical patterns due to the interactions of sound waves with surrounding buildings. An analysis of a rooftop installation of a heliport or vertiport provided some noise mitigation advantages. Noise contour distributions closely matched those outlined in previous research; however, the sound intensities modeled in this study were lower than indicated in exigent research. Assessments were also conducted for real-world sites with potential future use as vertiports in Seattle, Washington, and Chicago, Illinois. Practical Implications: This study suggests that current land use compatibility guidance may be inadequate for rotorcraft and eVTOL operations. Findings suggest the development of symmetrical noise-impact zones centered around heliports and vertiports. Examples of these zones are outlined, a, Peer Reviewed

Modern High Performance Computing (HPC) machines are distributed memoryclusters, consisting of multi-core compute nodes. Engineering simulation and analysis tools must employ efficient parallel algorithms in order to fully utilize the compute capability of modern HPC machines. The trend in Computational Fluid Dynamcis (CFD) has been to construct parallel solution algorithms based on some form of spatial domain decomposition. This approach has been shown to be a success for many practical applications. However, as one attempts to utlize more compute cores, limitations in strong scalability are inevitably reached due to a diminishing compute workload per compute core and either fixed or increasing communication cost. Furthermore, spatial domain decomposition approaches cannot be easily applied to mid-fidelity structural dynamics or rigid body dynamics models. A significant majority of industrial fluid and structural dynamic models utilize some form of time marching. Thus, if the domain decomposition strategy may be extended to include the temporal dimension, additional opportunity for increased parallelism may be realized. A new form of periodic multiple shooting is proposed that ismatrix-free and may be applied to high-fidelity multiphysics models or other high dimensional systems. The proposed methodology is formulated entirely in the time domain. Therefore, existing time-domain simulation tools may utilize the proposed approach to achieve a high degree of distributed memory parallelism without requiring any reformulation. Furthermore, the proposed methodology may be combined with conventional space domain decomposition techniques and other forms of data parallelism to achieve maximal performance on modern HPC architectures. The proposed algorithm retains the iterative shoot-correct approach of conventational periodic shooting methods. However, the correction stage is formulated using a hierarchical evaluation strategy combined with an Arnoldi subspace approximation

The idea of urban air mobility is steadily gaining popularity for its enormous potential to radically change the way modern cities are functioning and are being planned. In its present form, it relies on numerous vertical take-off and landing (VTOL) unmanned air vehicles (UAVs) of different sizes that will perform a number of tasks: deliveries, transport, surveillance and control, etc. On the other hand, to progress further, urban air mobility will also require abundant infrastructure on the ground (heliports, charging stations, communication hubs, etc.) as well as large amounts of energy that should be generated primarily from renewable resources. There are still many technological challenges to solve, such as efficient aircraft design, quiet rotors, electrical propulsion and storage, autonomous flight, optimized routes, etc. This paper describes an initial study of possible operation of several different UAVs (from the smallest to medium size) including the estimation of their required power and expected performance (such as useful weight, range and endurance) and proposes some options for future urban air mobility over Belgrade.

In this dissertation, a methodology is presented for lower-fidelity modeling of stacked rotors, with improved accuracy compared to existing lower-fidelity rotor models. This methodology is built on a conventional prescribed vortex wake model of the rotor, with lifting-line airfoil blades. Such a lifting-line model cannot fully capture the aerodynamic interaction between the rotor blades, which are driven heavily by thickness and shape effects. To account for these effects, a high-fidelity 2D CFD code is used to model the airfoil-to-airfoil interaction along the span of the rotor. These 2D CFD loads are then injected into the lifting-line/prescribed wake rotor model, using an iterative technique to account for the changing deflections of the rotor blades. To accurately determine the airfoil-interactional loads along the span of the rotor, it is necessary to have some way to relate the conditions along the span of the rotor in 3D to the airfoil conditions in 2D, and vice-versa. Methods for parameterization of the airfoil system are presented, which account for both the geometry of the rotor/airfoil system and their aerodynamic conditions. Two different methods of relating the airfoil loads back to the rotor are presented, which offer different strategies depending upon the constraints of the underlying rotor model. Any rotor design must include selection of the airfoils on the blades, and stacked rotors are no different. To that end, 2D airfoil simulations are presented, which demonstrate both the necessity of the current methodology, and offer suggestions for future stacked rotor design. 2D airfoil loads are pre-computed (prior to the lower-fidelity rotor simulations) using an established 2D CFD code. Automation of this code allows for rapid generation of large data sets with minimal user input. The combined 2D CFD/3D prescribed wake methodology is presented and validated against recent experimental results. The baseline prescribed wake model is shown to significantl

The research objective of this dissertation is to advance the understanding of rotorcraft broadband noise and reduce the noise impacts. Rotorcraft broadband noise has recently become a critical topic for vertical take-off and landing (VTOL) aircraft, due to the rapid progresses in Advanced Air Mobility (AAM) technologies. However, the current noise assessment tools of VTOL broadband noise rely heavily on semi-empirical models, which are short of both prediction accuracy and flow physics. To close this gap, a state-of-the-art prediction tool, namely UCD-QuietFly, is developed to assess rotor broadband noise using physics-based approaches. Extensive validations of UCD-QuietFly are performed against experiments. The effects of rotor design parameters on broadband noise are studied. Broadband noise impacts are investigated on Urban Air Mobility (UAM) conceptual designs, quiet helicopter designs, and small-scale drones. Finally, the noise reduction technique using trailing-edge serrations is analyzed. The first part of this dissertation investigates the effects of rotorcraft design and operating parameters on trailing-edge noise. A rotor trailing-edge noise prediction method is first developed. It is found that helicopter broadband noise scales with the 4.5th to 5.0th power of the tip Mach number in which the range is determined by the typical helicopter collective pitch angle in operation. Detailed trend analyses of noise levels as a function of frequency are presented in terms of the collective pitch angle, twist angle, rotor solidity, rotor radius, disk loading, and number of blades. Second, broadband noise of multi-rotor UAM vertical take-off and landing (VTOL) vehicles is studied. The multi-rotor broadband noise prediction capability is developed. It is found that UAM VTOL vehicles' broadband noise is important in the high-frequency range. For the same mission specifications, broadband noise is found to be higher for VTOL designs with more rotors. Multi-rotor vehicl

Airfoils have dominated the development of aircraft since the Wright brothers’ first flight. There are very few (if any) functional alternative mechanisms for heavier-than-air flight. However, in exploring the Magnus effect phenomenon applied to Flettner rotors in a rotorcraft configuration, a new and significantly underdeveloped method of heavier-than-air flight may be accomplished. Considering the aerodynamic context of the Magnus effect and its implementation in existing applications, this research principally concerns a single proposed mechanism and its design, viability, and lift-surface optimization. The proposed mechanism employs at least two 180o offset rotating cylinders rotating about a central vertical axis, like a helicopter rotor, with the backspin of the cylinders generating lift according to Bernoulli’s principle, Newton’s third law, and the Magnus effect.

In this research, a small-scale, 4-Kw, variable-voltage, hybrid-electric powertrain was constructed and tested to understand the fundamental behavior of such a system. The powertrain is meant for distributed propulsion for a multirotor electric vertical take-off and landing (eVTOL) aircraft. The powertrain was examined component by component, as well as in combination with one to four distributed rotors. Steady-state mathematical models for the engine, generator, and motor were developed for performance and weight. The component models were calibrated with the test data. It was found that a simple physics-based model of brake specific fuel consumption (bsfc) is possible to build if certain thermal efficiency constants could be calibrated with test data. The impact of various losses on the electric motor efficiency plots were revealed. Statistical weight models were developed by gathering a database of commercial reciprocating engines, electric motors, and power electronics, which were accurate to within ±30%. However, given the importance of electric motors for eVTOL design, a geometry- and material-based model was also developed. This model was accurate to within a 6% average error. The principal findings of this work are that the generator voltage is a key parameter in a hybrid-electic powertrain, and engine efficiency is closely coupled to the controls and aeromechanics of an eVTOL aircraft. The ability to vary the generator voltage with operating state appears crucial for the optimal specific fuel consumption. Generator voltage is a function of engine speed. For any operating state—defined by rotor torque and RPM—the generator voltage should be minimized as far as possible. However, generator voltage limits the maximum rotor RPM; so not all rotor RPM can be achieved at the same generator voltage. Hence, the optimal generator voltage will vary with rotor RPM as needed during specific mission segments. This implies for an RPM controlled aircraft, generator voltage

This work illustrates the use of artificial neural network modeling to study and characterize broadband blade-wake interaction noise from hovering sUAS rotors subject to varying airfoil geometries, rotor geometries, and operating conditions. Design of Experiments was used to create input feature spaces over 9 input features: the number of rotor blades, rotor size, rotor speed, the amount of blade twist, blade taper ratio, tip chord length, collective pitch, airfoil camber, and airfoil thickness. A high-fidelity strategy was then implemented at the discrete data points defined by the designed input feature spaces to design airfoils and rotor blades, predict the unsteady rotor aerodynamics and aeroacoustics, and isolate the blade-wake interaction noise from the acoustic broadband noise, which was then used for prediction model training and validation. An artificial neural network tool was developed and implemented into NASA's ANOPP2 code and was used to identify an optimal prediction model for the nonlinear functional relationship between the 9 input features and blade-wake interaction noise. This optimal artificial neural network was then validated over test data, and exhibited prediction accuracy over 91% for data previously unseen by the model. First- and second-order sensitivity analyses were then conducted using the developed artificial neural network tool and it was seen that input features which serve to directly modify the thrust coefficient, such as airfoil camber and collective pitch, had a dominant effect over blade-wake interaction noise, followed by second-order interaction effects related to the mean rotor solidity. The optimal prediction model along with aerodynamic simulations were used to further study the effect of varying input features on blade-wake interaction noise and three types of blade-wake interaction noise were identified. Blade-wake interaction noise caused by impingement of the turbulence entrained in a tip vortex on the leading edge of a

The aeromechanics of a slowed-rotor compound rotorcraft is investigated through wind-tunnel testing and comprehensive analysis. The emphasis is on a lift-offset wing compound with a hingeless rotor configuration. A new Maryland Compound Rig is developed and instrumented for wind-tunnel testing and an in-house rotor comprehensive code is modified and expanded for compound rotorcraft analysis. The compound rig consists of a lift compound model and a propeller model. The lift compound model consists of an interchangeable hub (articulated or hingeless), a fuselage, a half-wing of 70% rotor radius on the retreating side. The wing has a dedicated load cell and multiple attachment points relative to the rotor hub (16%R, 24%R, and 32%R and 5%R aft of the hub). The rotor diameter is 5.7-ft. The rotor has four blades with NACA 0012 airfoils with no twist and no taper. The wing incidence angle is variable between 0 to 12 degrees. The wing has a linearly varying thickness with symmetric airfoils NACA 0015 at the tip and NACA 0020 at the root. Sensors can measure rotor hub forces and moments, wing root forces and moments, blade pitch angles, structural loads (flap bending moment, lagbending moment, and torsional moment) at 25%R, pitch link loads, and hub vibratory loads. Wind tunnel tests are conducted up to advance ratio 0.7 for lift compound with half-wing at wing incidence angles of 4 and 8 degrees and compared with an isolated rotor. Hover tests are conducted up to tip Mach number of 0.5 to measure download penalty with the wing at various positions. The University of Maryland Advanced Rotorcraft Code (UMARC) is modified for compound rotorcraft analysis code. Aerodynamic models for the wing and the propeller are integrated. A recently developed Maryland Free Wake model is integrated, which can model the wake interaction between unequal and inharmonic speed rotor, wing, and propeller. The analysis is then validated with the test data. The validated analysis is used to analyze

The principal barrier of computational time for rotorcraft trim solution using high-fidelity three-dimensional (3D) structures on real rotor problems was overcome with parallel and scalable algorithms. These algorithms were devised by leveraging the modern supercomputer architecture. The resulting parallel X3D solver was used to investigate advanced coaxial rotors using a notional hingeless rotor test case, Metaltail. This investigation included rotor performance, blade airloads, vibratory hub loads, and three-dimensional stresses. The technical approach consisted of first studying existing algorithms for periodic rotor dynamics --- time marching, finite element in time (FET), and harmonic balance. The feasibility of these algorithms was studied for large-scale rotor structures, and drawbacks were identified. Modifications were then performed on the harmonic balance method to obtain a Modified Harmonic Balance (MHB) method. A parallel algorithm for skyline solver was devised on shared memory to obtain faster solutions to large linear system of equations. The MHB method was implemented on a hybrid distributed--shared memory architecture to allow for parallel computations of harmonics. These developed algorithms were then integrated into the X3D solver to obtain a new parallel X3D. The new parallel X3D was verified and validated in hover and forward flight conditions for both idealized and real rotor test cases. A total of four test cases were studied: 1) uniform beam, 2) Frank Harris rotor, 3) UH-60A-like Black Hawk rotor, and 4) NASA Tilt Rotor Aeroacoustic Model (TRAM). The predictions of tip displacements, airloads, and stress distributions from the MHB algorithm showed good agreement with the test data and time marching predictions. The key conclusion is that the new solver converges to the time marching solution 50-70 times faster and achieves a performance greater than 1 teraFLOPS. The new parallel X3D solver opened the opportunity for modeling advanced rotor c

Helicopters play a relevant role in society due to their extraordinary versatility. However, they are particularly vulnerable to adverse weather conditions as the majority of operations are carried out under visual flight rules. This is partly due to the shortage of tailored helicopter instrument flight procedures and routes. The emergence of the performance-based navigation concept supported by the latest satellite navigation technologies has opened up new possibilities for rotorcraft operations in the last few years. This paper presents an extensive overview of the state of the art in the design of performance-based navigation routes for helicopters from two main standpoints: instrument flight procedures and route spacing. Apart from summarizing recent and current major initiatives to implement helicopter low-level routes and flight procedures, this paper provides an outlook on the latest advances and ongoing efforts by the International Civil Aviation Organization in the field of helicopter procedure and route design to ensure flyability, obstacle clearance, strategical separation, and segregation of traffic flows. In addition, several gaps in the current design criteria are identified and suggestions for future research and development are outlined.

To ensure safe flight in icing conditions it is essential to understand ice accretion, its effect on aerodynamic performance and ice shedding in any aircraft certification program. Ice fragments shed from a rotorcraft windshield and other locations can be detrimental if they are ingested into the engine intake or impinge on the tail rotor. Because of the lack of experimental data about the forces and moments acting on ice in such a complex flow field, the computational simulation of ice shedding trajectories becomes essential. This study presents a methodology to predict the location of ice accretion and break-off, ice shape and shedding trajectory in a rotorcraft flow field with strong rotor wakes during forward flight. The methodology includes the creation of an aerodynamic database for different ice shapes (rectangle, disc, ellipse, and glaze ice shape) at various combinations of Euler angles; the analysis of rotorcraft flow field by computational fluid dynamics for different advance ratios; and a six degree-of-freedom trajectory analysis using artificial neural networks and the Monte Carlo method. The actuator surface method was applied to account for the rotor wake effect, which is capable of modeling the tip vortices and inboard sheets emanating from the rotor blades. The main results are a probability map of the ice shedding trajectory footprints on the engine intake and tail rotor planes. Disc-shaped ice fragments with a sharp edge turned out to be most dangerous. The rotation of the main rotors also substantially affected ice accretion and shed trajectory, indicating the importance of integrated simulations of all components when designing ice protection systems for rotorcraft.

Search and rescue missions using rotorcrafts need to be reliable all year long, even in winter conditions. In some cases of deployment prior to take off, the crew may need to manually remove accumulated contaminant from the critical surfaces using tools at their disposal. However, icy contaminant may be hard to remove since the rotorcrafts critical surfaces could be cooler than the environment, thus promoting adhesion. Currently, there exists several passive ice protection materials that could reduce the ice adhesion strength and assist the manual de-icing. The aim of this paper is to propose a detailed comparative procedure to assess the ability of materials to assist the manual de-icing of rotorcrafts. The proposed procedure consists of the characterization of materials using several laboratory tests in order to determine their characteristics pertaining to wettability, their icephobic behavior, and finally their assessment under a multi-tool analysis to evaluate if they can assist. The multi-tool analysis uses different mechanical tools, which are currently used during normal operation, to execute a gradual de-icing procedure, which begins with the softest to the hardest tool using a constant number of passes or strokes, under different types of simulated precipitation. Five different materials were used to evaluate the proposed procedure: Aluminum (used as a reference), two silicone-based coatings (Nusil and SurfEllent), an epoxy-based coating (Wearlon), and finally a commercial ski wax (Swix). All of the tested materials could assist the manual de-icing, within a certain limit, when compared to the bare aluminum. However, SurfEllent was the material that obtained the best overall results. This procedure could be easily adapted to different fields of application and could be used as a development tool for the optimization and the assessment of new materials aimed to reduce ice adhesion.

In order to study ice protection systems for rotating blades, a new experimental setup has been developed at the Anti-Icing Materials International Laboratory (AMIL). This system consists of two small-scale rotating blades in a refrigerated icing wind tunnel where atmospheric icing can be simulated. Power is brought to the blades through a slip ring, through which the signals of the different sensors that are installed on the blades also pass. As demonstrated by the literature review, this new setup will address the need of small-scale wind tunnel testing on electrically powered rotating blades. To test the newly designed apparatus, preliminary experimentation is done on a hybrid ice protection system. Electrothermal protection is combined with different surface coatings to measure the impact of those coatings on the power consumption of the system. In anti-icing mode, the coatings tested did not reduce the power consumption on the system required to prevent ice from accumulating on the leading edge. The coatings however, due to their hydrophobic/superhydrophobic nature, reduced the power required to prevent runback ice accumulation when the leading edge was protected. One of the coatings did not allow any runback accumulation, limiting the power to protect the whole blades to the power required to protect solely the leading edge, resulting in a potential 40% power reduction for the power consumption of the system. In de-icing mode, the results with all the substrates tested showed similar power to achieve ice shedding from the blade. Since the coatings tested have a low icephobicity, it would be interesting to perform additional testing with icephobic coatings. Also, a small unheated zone at the root of the blade prevented complete ice shedding from the blade. A small part of the ice layer was left on the blade after testing, meaning that a cohesive break had to occur within the ice layer, and therefore impacting the results. Improvements to the setup will be done

Aircraft in-flight icing is an important problematic in the aerospace industry, especially for small helicopters. All rotorcraft currently habilitated to fly under icing conditions are equipped with electrothermal systems. They are not fitted to small and even medium size helicopters, which cannot provide the high power required and bear the additional weight of those systems. As an alternative, this research investigates the use of piezoelectric actuators to develop a low-energy vibration based ice protection system that could be implemented on small rotorcraft. The objective of this project was divided in four parts: (1) to design a piezoelectric actuator-based de-icing system integrated to a flat plate experimental setup and develop a numerical model of the system with experimental validation, (2) use the experimental setup to investigate actuator activation with frequency sweeps and transient vibration analysis, (3) add ice layer to the numerical model and predict numerically stresses for different ice breaking with experimental validation, and (4) bring the concept to a blade structure for wind tunnel testing. First, preliminary numerical analysis was performed to gain basic guidelines for the integration of piezoelectric actuators in a simple flat plate experimental setup for vibration-based de-icing investigation. The results of these simulations allowed to optimize the positioning of the actuators on the structure and the optimal phasing of the actuators for mode activation. A numerical model of the final setup was elaborated and an experimental setup was fabricated faithful to the numerical model at the laboratory with piezoelectric actuator patches bonded to a steel flat plate. The experimental setup was brought at the National Research Council Canada (NRC) for testing with a laser vibrometer to validate the numerical results. The experimental results validated the model when the plate is optimally excited with an average of error of 20% and a maximal erro

Calculating the unsteady convective heat transfer on helicopter blades is the first step in the prediction of ice accretion and the design of ice-protection systems. Simulations using Computational Fluid Dynamics (CFD) successfully model the complex aerodynamics of rotors as well as the heat transfer on blade surfaces, but for a conceptual design, faster calculation methods may be favorable. In the recent literature, classical methods such as the blade element momentum theory (BEMT) and the unsteady vortex lattice method (UVLM) were used to produce higher fidelity aerodynamic results by coupling them to viscous CFD databases. The novelty of this research originates from the introduction of an added layer of the coupling technique to predict rotor blade heat transfer using the BEMT and UVLM. The new approach implements the viscous coupling of the two methods from one hand and introduces a link to a new airfoil CFD-determined heat transfer correlation. This way, the convective heat transfer on ice-clean rotor blades is estimated while benefiting from the viscous extension of the BEMT and UVLM. The CFD heat transfer prediction is verified using existing correlations for a flat plate test case. Thrust predictions by the implemented UVLM and BEMT agree within 2% and 80% compared to experimental data. Tip vortex locations by the UVLM are predicted within 90% but fail in extreme ground effect. The end results present as an estimate of the heat transfer for a typical lightweight helicopter tail rotor for four test cases in hover, ground effect, axial, and forward flight.

The objective of this research project is divided in four parts: (1) to design a piezoelectric actuator-based de-icing system integrated to a flat plate experimental setup and develop a numerical model of the system with experimental validation, (2) use the experimental setup to investigate actuator activation with frequency sweeps and transient vibration analysis, (3) add an ice layer to the numerical model and predict numerically stresses at ice breaking with experimental validation, and (4) bring the concept to a blade structure for wind tunnel testing. This paper presents the third part of the investigation in which an ice layer is added to the numerical model. Five accelerometers are installed on the flat plate to measure acceleration. Validation of the vibration amplitude predicted by the model is performed experimentally and the stresses calculated by the numerical model at cracking and delamination of the ice layer are determined. A stress limit criteria is then defined from those values for both normal stress at cracking and shear stress at delamination. As a proof of concept, the numerical model is then used to find resonant modes susceptible to generating cracking or delamination of the ice layer within the voltage limit of the piezoelectric actuators. The model also predicts a voltage range within which the ice breaking occurs. The experimental setup is used to validate positively the prediction of the numerical model.

The objective of this research project is divided in four parts: (1) to design a piezoelectric actuator-based de-icing system integrated to a flat plate experimental setup and develop a numerical model of the system with experimental validation, (2) use the experimental setup to investigate actuator activation with frequency sweeps and transient vibration analysis, (3) add ice layer to the numerical model and predict numerically stresses for different ice breaking with experimental validation, and (4) bring the concept to a blade structure for wind tunnel testing. This paper presents the first objective of this study. First, preliminary numerical analysis was performed to gain basic guidelines for the integration of piezoelectric actuators in a simple flat plate experimental setup for vibration-based de-icing investigation. The results of these simulations allowed to optimize the positioning of the actuators on the structure and the optimal phasing of the actuators for mode activation. A numerical model of the final setup was elaborated with the piezoelectric actuators optimally positioned on the plate and meshed with piezoelectric elements. A frequency analysis was performed to predict resonant frequencies and mode shapes, and multiple direct steady-state dynamic analyses were performed to predict displacements of the flat plate when excited with the actuators. In those steady-state dynamic analysis, electrical boundary conditions were applied to the actuators to excite the vibration of the plate. The setup was fabricated faithful to the numerical model at the laboratory with piezoelectric actuator patches bonded to a steel flat plate and large solid blocks used to mimic perfect clamped boundary condition. The experimental setup was brought at the National Research Council Canada (NRC) for testing with a laser vibrometer to validate the numerical results. The experimental results validated the model when the plate is optimally excited with an average of error of 20

The objective of this research project is divided in four parts: (1) to design a piezoelectric actuator based de-icing system integrated to a flat plate experimental setup, develop a numerical model of the system and validate experimentally; (2) use the experimental setup to investigate actuator activation with frequency sweeps and transient vibration analysis; (3) add an ice layer to the numerical model, predict numerically stresses at ice breaking and validate experimentally; and (4) implement the concept to a blade structure for wind tunnel testing. This paper presents the second objective of this study, in which the experimental setup designed in the first phase of the project is used to study transient vibration occurring during frequency sweeps. Acceleration during different frequency sweeps was measured with an accelerometer on the flat plate setup. The results obtained showed that the vibration pattern was the same for the different sweep rate (in Hz/s) tested for a same sweep range. However, the amplitude of each resonant mode increased with a sweep rate decrease. Investigation of frequency sweeps performed around different resonant modes showed that as the frequency sweep rate tends towards zero, the amplitude of the mode tends toward the steady-state excitation amplitude value. Since no other transient effects were observed, this signifies that steady-state activation is the optimal excitation for a resonant mode. To validate this hypothesis, the flat plate was installed in a cold room where ice layers were accumulated. Frequency sweeps at high voltage were performed and a camera was used to record multiple pictures per second to determine the frequencies where breaking of the ice occur. Consequently, the resonant frequencies were determined from the transfer functions measured with the accelerometer versus the signal of excitation. Additional tests were performed in steady-state activation at those frequencies and the same breaking of the ice layer was o

Flow separation is prevalent in a number of aerospace applications, but because of complex non-linearities in the governing equations, the resulting aerodynamic forces are challenging to model. This mathematical limitation is particularly impactful in the field of high speed rotorcraft. When a rotor operates at high advance ratio, a regime typical of high speed (and low power) flight, the blades of the rotor are subject to several unsteady motions that incur flow separation, including high pitch inputs and a region of reverse flow that occupies much of the rotor's retreating side. The aerodynamic forces in these regions are dominated by large-scale, coherent vortex structures that are poorly captured by conventional aerodynamics theories. The purpose of this work is to understand the physics of flow separation on high advance ratio rotors, and to leverage this understanding into a low-order, physics-based model for use in rotorcraft design applications. The current work approaches this goal by identifying, understanding, and ultimately modeling the coherent flow structures present on a representative, sub-scale rotor system operating at high advance ratio. Flowfield measurements on this rotor revealed the presence of two distinct vortex structures, a ``sharp-edge'' vortex and a ``blunt-edge'' vortex, believed to dominate unsteady loading in the reverse flow region. The sharp-edge vortex was studied via a high-fidelity numerical simulation, and its growth was found to be dominated by 2-D mechanisms of vorticity transport. The insignificance of 3-D effects was attributed to a mutual cancellation of Coriolis forces and spanwise convection/tilting, a feature unique to reverse flow. Likewise, the blunt-edge vortex was studied in a series of 2-D surging and pitching wing experiments; its formation was found to largely depend on the unsteady, "external" features of the flow, most notably the trailing wake. Together, these observations led to the development of a 2-D discre

In forward flight, slowing down a rotor alleviates compressibility effects on the advancing side, extending the cruise speed limit and inducing high advance ratio flight regime. To investigate the aerodynamic phenomena at high advance ratios and provide data for the validation of analysis tools, a series of wind tunnel tests were conducted progressively with a 33.5-in radius, 4-bladed Mach-scaled rotor in the Glenn L. Martin Wind Tunnel. In the first stage of the research, a wind tunnel test was carried out at high advance ratios with highly similar, non-instrumented blades and on-hub control angle measurements, in order to gain a baseline performance and control dataset with minimum error due to blade structural dissimilarity and pitch angle discrepancy. The tests were conducted at advance ratios of 0.3 to 0.9, and a parametric study on shaft tilt was conducted at $0^\circ$ and $\pm 4^\circ$ shaft tilt angles. The test data were then compared with those of previous tests and with the predictions of the in-house comprehensive analysis UMARC. The airload results were investigated using comprehensive analysis to gain insights on the influences of advance ratio and shaft tilt angle on rotor performance and hub vibratory loads. Results indicate that the thrust benefit from backward shaft tilt is dependent on the change in the inflow condition and the induced angle of attack increment, and the reverse flow region at high advance ratios is the major contributor to changes in shaft torque and horizontal force. In the second stage of the research, the rotor blades were instrumented with pressure sensors and strain gauges at 30\% radius, and pressure data were acquired to calculate the sectional airloads by surface integration up to an advance ratio of 0.8. The test results of blade airloads and structural loads were compared with the predictions of comprehensive analysis (UMARC and PrasadUM) and CFD/CSD coupled analysis (PrasadUM/HAMSTR). The focus was on the data correlati

A multi-solver, overset, computational fluid dynamics framework is developed for efficient, large-scale simulation of rotorcraft problems. Two primary features distinguish the developed framework from the current state of the art. First, the framework is designed for heterogeneous compute architectures, making use of both traditional codes run on the Central Processing Unit (CPU) as well as codes run on the Graphics Processing Unit (GPU). Second, a framework-level implementation of the Generalized Minimal Residual linear solver is used to consider all meshes from all solvers in a single linear system. The developed GPU flow solver and framework are validated against conventional implementations, achieving a 5.35× speedup for a single GPU compared to 24 CPU cores. Similarly, the overset linear solver is compared to traditional techniques, demonstrating the same convergence order can be achieved using as few as half the number of iterations. Applications of the developed methods are organized into two chapters. First, the heterogeneous, overset framework is applied to a notional helicopter configuration based on the ROBIN wind tunnel experiments. A tail rotor and hub are added to create a challenging case representative of a realistic, full-rotorcraft simulation. Interactional aerodynamics between the different components are reviewed in detail. The second application chapter focuses on performance of the overset linear solver for unsteady applications. The GPU solver is used along with an unstructured code to simulate laminar flow over a sphere as well as laminar coaxial rotors designed for a Mars helicopter. In all results, the overset linear solver out-performs the traditional, de-coupled approach. Conclusions drawn from both the full-rotorcraft and overset linear solver simulations can have a significant impact on improving modeling of complex rotorcraft aerodynamics.

The objective of this thesis is a comprehensive investigation of hydrogen fuel cells for electric vertical take-off and landing (eVTOL) aircraft. The primary drawback of battery powered eVTOL aircraft is their poor range and endurance with practical payloads. This work uses simulation and hardware testing to examine the potential of hydrogen fuel cells to overcome this drawback. The thesis develops steady state and transient models of fuel cells and batteries, and validates the models experimentally. An equivalent circuit network model was able to capture the waveforms and magnitudes of voltage as a function of current. Temperature and humidity corrections were also included. Examination of the results revealed that the transient behavior of batteries and fuel stacks are significant primarily shortly after startup of the fuel stack and at the limiting ranges of high and low power; for a nominal operating power and barring faults, steady state models were adequate. This work then demonstrates fuel cell and battery power sharing in regulated and unregulated parallel configurations. It details the development of a regulated architecture, which controls power sharing, to achieve a reduction in power plant weight. Finally, the thesis outlines weight models of motors, batteries, and fuel cells needed for eVTOL sizing, and carries out sizing analysis for on-demand urban air taxi missions of three different distances -- 50, 75, and 150~mi of cruise and 5~min total hover time. This revealed that for ranges within 75 mi, a light weight (5000-6000~lb gross weight) all-electric tilting proprotor configuration achieves a practical payload (500~lb or more) with current levels of battery specific energy (150~Wh/kg) if high burst C-rate batteries are available (4-10~C for 2.5~min). Either a battery-only or battery-fuel cell (B-FC) hybrid power plant is ideal depending on the range of the mission: For inter-city ranges (beyond approximately 50~mi), the mission is impossible with bat

Modern aerodynamic analysis tools, such as free-vortex wake models and CFD-based techniques, include fewer theoretical limitations and approximations than classical simplified schemes, and represent the state-of-the-art in rotorcraft aerodynamic modeling, including for coaxial and other advanced configurations. However, they are impractical or impossible to apply to many flight dynamics problems because they are not formulated in ordinary differential equation (ODE) form, and they are often computationally intensive. Inflow models, for any configuration type, that couple the accuracy of high-fidelity aerodynamic models with the simplicity and ODE form of dynamic inflow-type theories would be an important contribution to the field of flight dynamics and control. This dissertation presents the methodology for the extraction of linearized ODE models from computed inflow data acquired from detailed aerodynamic free-vortex wake models, using frequency domain system identification. These methods are very general and applicable to any aerodynamic model, and are first demonstrated with a free wake model in hover and forward flight, for a single main rotor, and subsequently for the prediction of induced flow off the rotor as well, at locations such as the tail or fuselage. The methods are then applied to the extraction of first order linearized ODE inflow models for a coaxial rotor in hover. Subsequent analysis concluded that free-vortex wake models show that the behavior of the inflow of a coaxial configuration may be higher-order. Also, tip-path plane motion of a coaxial rotor causes wake distortion which has an impact on the inflow behavior. Therefore, the methodology is expanded to the identification of a second order inflow representation which is shown to better capture from all of the relevant dynamics from free-vortex wake models, including wake distortion. With ODE models of inflow defined for an advanced coaxial configuration, this dissertation then presents a comp

Flight safety has been a fundamental aspect of aircraft, and the future demand for wider usage of aerial operations leads to more focus on the flight safety. Particularly rotorcraft require high standards of flight safety due to their inherent features, such as complicated rotary mechanisms, close-to-ground operations, and complex aerodynamic environment. Consequently, rotorcraft pilots need to exert relatively high workload to safely operate these vehicles. An understanding of the interaction between the rotorcraft and the pilot is essential for improving flight safety. This interaction is elaborated by the Handling Qualities (HQ) discipline, which aims to identify and, if possible predict any deficiency in HQ that could potentially jeopardize safe flight. A typical (and potentially catastrophic) example of a HQ deficiency are the Aircraft / Rotorcraft Pilot Couplings (A/RPC), formerly referred to as Pilot Induced Oscillations (PIO). A/RPC is defined as the involuntary and adverse interaction between the pilot and the vehicle under control. Generally for rotorcraft, the ‘vehicle’ part of this interaction is evaluated by objective HQ criteria and online Rotorcraft Pilot Coupling (RPC) detection tools, whereas the ‘pilot’ part is assessed with subjective pilot ratings. Using subjective ratings has several disadvantages, such as being used at very late stages of the design when a prototype vehicle is already built. Addressing a serious HQ deficiency after this late design stage then requires immerse effort to re-design the vehicle systems and repeat the flight tests..., Control & Simulation

The objective of this paper is to raise awareness about the significance of collecting and ensuring the quality of the obstacle data required for the safety of air navigation for both manned and unmanned aviation. This information could be of importance to geodetic, CityGML, 3D model and Building Information Management (BIM) community. With the advancement of future air mobility concepts such as drones and Personal Air Vehicles (PAVs), there is an increased demand for obstacle data of higher accuracy, including at Very Low Level (VLL) altitude. The paper presents the requirements pertaining to aviation such that the above mentioned communities could understand the existing complexity. This complexity adheres to the aggregation of quality-assured obstacle data from domains outside the aviation field’s responsibility. It is expected that with model developments (e.g. BIM), new solutions could be identified to support the aviation community with the aggregation of obstacle data of required quality., Control & Simulation

Aircraft manufacturers are increasingly exploring emission-free flight or emission reduction for larger passenger aircraft. The low energy density of state-of-the-art battery technology limits the application to small, electric, fixed wing aircraft up to a flight time of approximately one hour. To overcome these limits, a combination of fuel cells and batteries to exploit the benefits of battery power density and hydrogen energy density was studied. Current Lithium-Ion battery cells reach approximately 1.6 kW/kg of maximum power density, much higher than fuel cell systems. On the other hand, the energy storage capacity of suitable hydrogen storage methods is much larger than battery cells, the latter have an energy density of 240 Wh/kg. Because most demonstrated applications are for fixed wing aircraft, the unmanned GeoCopter GC-201 helicopter was used for performance requirements, weight and volume analysis. The study focuses on the preliminary sizing of the powertrain and the optimization of fuel cell and mission profile variables for this vehicle. Helicopter performance modelling, fuel cell static behavior as well as a battery discharge simulation are combined with lower fidelity models for other components. The study results in a comparison of battery-only and fuel cell-battery configurations through payload-range diagrams, allowing for a quick evaluation of application areas. These mainly show that batteries excel at high payload, low range applications whereas a fuel cell-battery combination shows clear advantages at low payload, longer range applications. Liquid hydrogen will be shown to be comparable to the current micro gas turbine powered rotorcraft, with 400 and 500 km range capabilities respectively. Range capabilities for 300 bar and 700 bar compressed gas tank storage options show 140 and 180 km, with battery-only reaching a maximum range of 80 km., Aerospace Engineering

Rotorcraft flyover noise has long been a field of study for researchers. This is because for many people, the sounds produced by these vehicles are found to be extremely annoying. The focus of this thesis is to recreate the time-varying rotorcraft noise at the source for a single emission angle. Then, through interpolation between emission angles, produce a simulated flyover at the source that can then be propagated to a receiver. This will allow for the creation of a simulated flyover without the need of having to use a physical aircraft, or pre-existing data from some type of data collection means, such as a microphone array. The current methods are limited to a predefined length of data in order to synthesize signals. It has been documented that synthesizing flyover noise, from direct use of physical flyover recordings through an empirical approach, yields a high fidelity signal, as long as both unmodulated and modulated components are present. In order to extend these signals indefinitely, models for the amplitude and phase modulation must be developed. A band-limited random process will be explored for both the amplitude and phase modulations. An overlap-add technique, as well as a randomization technique and a modified phase modulation signal, defined as the "residual", will also be attempted in order to model the phase modulation. The results from this work have successfully found a means in which to produce a viable model of the amplitude modulation. Further investigation is still required in order to produce a model of the phase modulation which results in a high-fidelity model that can be extended indefinitely.

This dissertation describes the first comprehensive use of integrated 3-D aeromechanics modeling, defined as the coupling of 3-D solid finite element method (FEM) structural dynamics with 3-D computational fluid dynamics (CFD), for the analysis of a real helicopter rotor. The development of this new methodology (a departure from how rotor aeroelastic analysis has been performed for 40 years), its execution on a real rotor, and the fundamental understanding of aeromechanics gained from it, are the key contributions of this dissertation. This work also presents the first CFD/CSD analysis of a tiltrotor in edgewise flight, revealing many of its unique loading mechanisms. The use of 3-D FEM, integrated with a trim solver and aerodynamics modeling, has the potential to enhance the design of advanced rotors by overcoming fundamental limitations of current generation beam-based analysis tools and offering integrated internal dynamic stress and strain predictions for design. Two primary goals drove this research effort: 1) developing a methodology to create 3-D CAD-based brick finite element models of rotors including multibody joints, controls, and aerodynamic interfaces, and 2) refining X3D, the US Army’s next generation rotor structural dynamics solver featuring 3-D FEM within a multibody formulation with integrated aerodynamics, to model a tiltrotor in the edgewise conversion flight regime, which drives critical proprotor structural loads. Prior tiltrotor analysis has primarily focused on hover aerodynamics with rigid blades or forward flight whirl-flutter stability with simplified aerodynamics. The first goal was met with the development of a detailed methodology for generating multibody 3-D structural models, starting from CAD geometry, continuing to higher-order hexahedral finite element meshing, to final assembly of the multibody model by creating joints, assigning material properties, and defining the aerodynamic interface. Several levels of verification and valida

The purpose of this research is to further the understanding of rotor aeromechanics at advance ratios (mu) beyond the maximum of 0.5 (ratio of forward airspeed to rotor tip speed) for conventional helicopters. High advance ratio rotors have applications in high speed compound helicopters. In addition to one or more conventional main rotors, these aircraft employ either thrust compounding (propellers), lift compounding (fixed-wings), or both. An articulated 4-bladed model rotor was constructed, instrumented, and tested up to a maximum advance ratio of mu=1.6 in the Glenn L. Martin Wind Tunnel at the University of Maryland. The data set includes steady and unsteady rotor hub forces and moments, blade structural loads, blade flapping angles, swashplate control angles, and unsteady blade pressures. A collective-thrust control reversal---where increasing collective pitch results in lower rotor thrust---was observed and is a unique phenomenon to the high advance ratio flight regime. The thrust reversal is explained in a physical manner as well as through an analytical formulation. The requirements for the occurrence of the thrust reversal are enumerated. The effects of rotor geometry design on the thrust reversal onset are explored through the formulation and compared to the measured data. Reverse-flow dynamic stall was observed to extend the the lifting capability of the edgewise rotor well beyond the expected static stall behavior of the airfoil sections. Through embedded unsteady blade surface pressure transducers, the normal force, pitching moment, and shed dynamic stall vortex time histories at a blade section in strong reverse flow were analyzed. Favorable comparisons with published 2-D pitching airfoil reverse flow dynamic stall data indicate that the 3-D stall environment can likely be predicted using models developed from such 2-D experiments. Vibratory hub loads were observed to increase with advance ratio. Maximum amplitude was observed near mu=1, with a reduct

The constant need to improve helicopter performance requires the optimization of existing and future rotor designs. A crucial indicator of rotor capability is hover performance, which depends on the near-body flow as well as the structure and strength of the tip vortices formed at the trailing edge of the blades. Computational Fluid Dynamics (CFD) solvers must balance computational expenses with preservation of the flow, and to limit computational expenses the mesh is often coarsened in the outer regions of the computational domain. This can lead to degradation of the vortex structures which compose the rotor wake. The current work conducts three-dimensional simulations using OVERTURNS, a three-dimensional structured grid solver that models the flow field using the Reynolds-Averaged Navier-Stokes equations. The S-76 rotor in hover was chosen as the test case for evaluating the OVERTURNS solver, focusing on methods to better preserve the rotor wake. Using the hover condition, various computational domains, spatial schemes, and boundary conditions were tested. Furthermore, a mesh adaption routine was implemented, allowing for the increased refinement of the mesh in areas of turbulent flow without the need to add points to the mesh. The adapted mesh was employed to conduct a sweep of collective pitch angles, comparing the resolved wake and integrated forces to existing computational and experimental results. The integrated thrust values saw very close agreement across all tested pitch angles, while the power was slightly over predicted, resulting in under prediction of the Figure of Merit. Meanwhile, the tip vortices have been preserved for multiple blade passages, indicating an improvement in vortex preservation when compared with previous work. Finally, further results from a single collective pitch case were presented to provide a more complete picture of the solver results.

Consulta en la Biblioteca ETSI Industriales (Riunet), [ES] Los avances en el campo de la robótica, en particular el de los sistemas aéreos no tripulados (UAV), han tenido un desarrollo explosivo en las últimas décadas. Gran parte de estas investigaciones han sido financiadas gracias a fondos militares, que han hallado en estas plataformas una solución óptima para llevar a cabo tareas de inteligencia y vigilancia (ISAR). Las ventajas principales que aportan los UAV consisten en la posibilidad de operar en ambientes hostiles en cualquier aspecto, así como facilitar la realización de misiones de muy larga duración con cierto grado de monotonía. No obstante, la posibilidad de aplicar estas tecnologías en tareas civiles ha conseguido aumentar el interés de la industria por estos sistemas. Entre las aplicaciones potenciales pueden encontrarse la realización de fotografía aérea para el mapeado de terrenos, levantamientos topográficos, tareas de vigilancia urbana, búsqueda seguimiento e inspección en entornos urbanos e industriales, exploración de edificios y túneles, etc. De entre las posibles configuraciones, los vehículos basados en rotor para generar sustentación se caracterizan por suponer un mayor reto de control por su inestabilidad inherente. A pesar de ello, su capacidad para mantener una posición fija en el espacio así como su maniobrabilidad las convierten en plataformas idóneas para tareas de búsqueda y rescate. Sin embargo, para poder asegurar un grado aceptable de rendimiento en el desempeño de la misión, es necesario haber desarrollado el correspondiente grado de especialización de la plataforma. Para ello es requisito indispensable adquirir un conocimiento detallado tanto de la plataforma como del entorno de operación., [EN] This thesis is part of the project “Design of a spherical UAV for operations in complex environments”. The cited project is divided in two parts, the one corresponding to this thesis “Modelling and simulation” and the one developed by Roland Dixon in “Design and control” (Dixon 2013). The main contribution of this thesis towards the global aim of the complete design and manufacturing of a prototype Unmanned Aerial Vehicle (UAV) is the mathematical modelling of the system. Subsequently this research also includes the development of the tools and techniques needed to identify the aforementioned UAV. As both the model and the prototype syntheses are undertaken in parallel, the achievements and discoveries in one are fed back to enhance the other. Consequently two models were developed. One defines the flaps effect as that of an aerofoil and, thus, is based on the identification of the lift and drag coefficients functions. The other understands the vehicle propulsion system as a nozzle with thrust vectoring fins interposed in the outlet. The comparison of the gathered data with the simulation outputs clearly shows that thrust vectoring is the best modelling of the behaviour of the vehicle. The dynamics of yaw are identified as a first order system with an added zero and a time constant of 1.12. Moreover, both pitch and roll dynamics can be expressed as a second order system with a natural frequency of 0.707 and a damping coefficient of 19.54. The early versions of the model were successfully used for the synthesis of a robust controller that enabled flight testing. Finally, the model is validated using this data and verifying the real system dynamics.

The aeroacoustic research activity of rotorcraft in Japan has been discussed. The main institutes, universities and companies have participated in the panel discussion, namely Tokyo University, National Aerospace Laboratory, Advanced Technology Institute of Commuter Helicopter (ATIC), Ishikawajima-Harima, Kawasaki, Mitsubishi and Fuji Heavy Industries. The topics to discuss are as follows: the brief talk with the aeroacoustic research activity in each institutes and companies, the measurement technology and its application in each institutes and companies, the present status of aeroacoustic analysis including CFD (Computational Fluid Dynamics) techniques and its application into design, the research activity for rotorcraft aeroacoustics in NAL and its role for other institutes and companies and the research cooperation and its organization with foreign research institutes. This report describes the brief summary of the panel discussion., 日本における回転翼機の空力騒音研究についてパネルディスカッションの形式で討論が行われた。パネリストは東京大学、航空宇宙技術研究所、コミュータヘリコプタ先進技術研究所(ATIC)、石川島播磨重工業、川崎重工業、三菱重工業、富士重工業の主な代表者である。議論された内容は、各機関・各社における回転翼騒音研究の概要(機内、機外)、ヘリコプタ騒音の計測技術と対策、CFD(数値流体力学)を含む騒音解析の現状と設計への反映、航技研での回転翼騒音研究の役割と航技研への要望、国内・国外との研究協力および体制である。本報告は、パネルディスカッションでの討論内容を要約したものである。

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 aspects 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 with using blade planform shapes and active blade control techniques. The important parameters for rotor blade-vortex interaction noise and vibration are identified: blades tip vortex structures and its trajectory, blade aeroelastic deformation, and airloads. The current prediction capabilities of rotor noise are mature with the known blade airloads. But the current prediction capabilities of blade airloads and rotor wakes have serious shortcomings. Many blade tip design concepts have been investigated for diffusing tip vortices and eventually for reducing noise. However, these tip shapes have not been able to substantially reduce blade-vortex interaction noise. Higher harmonic pitch control and individual blade control concepts have been extensively investigated for noise and vibration reduction in recent years. The higher harmonic pitch control technique has proven the substantial noise reduction, up to 6 dB, while vibration and low frequency noise have been increased. Individual blade control tests have shown the simultaneous reduction of rotor noise and vibratory loads with 2 pitch/rev control inputs. Recently, active blade control concepts with smart structures have been actively pursued 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., ヘリコプタの主回転翼によって発生する回転翼と渦の相互作用による騒音は最も重要な騒音問題の1つであり、軍事的な観点からと同時にヘリコプタの地域社会での受け入れとの面からも非常に重要である。数10年にわたる研究の結果、騒音発生のメカニズムに対する物理的な理解は実質的に改善され、回転翼平面の形状およびアクティブ回転翼制御技術を用い、騒音の放射を制御するために多様な設計概念が研究されて来た。回転翼と渦の相互作用による騒音と振動に対する重要なパラメータが特定された:回転翼端渦の形状、その軌跡、回転翼の空力弾性変形および空気抗力。回転翼の空気抗力が既知であると、現在、回転翼騒音を予想する能力は成熟している。しかし翼の空気抗力抵抗および回転翼の後流を予想する現状の能力不足は深刻である。翼先端の渦を拡散し結果として騒音を低減するために、翼先端の設計概念について多くの研究がなされて来た。しかしこれら先端形状により翼と渦の相互作用による騒音を実質的に低減をする事が出来なかった。近年騒音と振動を低減するため、高調波ピッチの制御と個々の翼制御に対する概念が大規模に研究されて来た。高調波ピッチの制御技術により6dBまで実質的に騒音を低減できることが証明されたが、一方、振動と低周波騒音は増加した。個々のブレード制御試験により1回転2ピッチの制御入力により回転翼騒音と振動荷重を同時に低減出来ることが証明された。最近、アクティブ翼捻りや後縁のフラップを重視し、スマート構造を持ったアクティブ制御概念が積極的に追及されている。スマート構造技術は非常に有望であるが、回転翼機の応用で、振動数、力、変位などの全ての要求に対応するためには更なる進歩が必要である。

This dissertation presents a lifting-line, comprehensive approach to predicting the performance and loads of high advance ratio rotorcraft. At high advance ratios, the reverse flow region is large and its unique aerodynamics impacts the rotor performance and dynamics more than at conventional airspeeds where they are often ignored. The analysis is refined and augmented with improved modeling of the nearwake in reverse flow, a new aerodynamic model of the fuselage and the root cutout region and corrections to the airfoil properties for highly yawed flow. The analysis is correlated and evaluated against a full-scale UH-60A rotor test to an advance ratio of 1.0 and against an in-house Mach-scaled rotor to an advance ratio of 1.2. High advance ratio performance is predicted satisfactorily for both tests, including predicting the onset of thrust reversal. Despite the high advance ratio, correctly modeling the wake is most important for predicting airloads and the resulting blade bending loads, while yawed flow, nearwake inflow and the fuselage flow disturbances are important for predicting high advance ratio thrust and power. The validated analysis is used to investigate the effect of reverse flow stall, blade twist, root cut-out and shaft angle on high advance ratio performance.

The focus of this research was to demonstrate a four blade rotor trim in forward flight using integrated trailing edge flaps instead of using a swashplate controls. A compact brushless DC motor was evaluated as an on-blade actuator, with the possibility of achieving large trailing edge flap amplitudes. A control strategy to actuate the trailing edge flap at desired frequency and amplitude was developed and large trailing edge flap amplitudes from the motor (instead of rotational motion) were obtained. Once the actuator was tested on the bench-top, a lightweight mechanism was designed to incorporate the motor in the blade and actuate the trailing edge flaps. A six feet diameter, four bladed composite rotor with motor-flap system integrated into the NACA 0012 airfoil section was fabricated. Systematic testing was carried out for a range of load conditions, first in the vacuum chamber followed by hover tests. Large trailing edge flap deflections were observed during the hover testing, and a peak to peak trailing edge flap amplitude of 18 degree was achieved at 2000 rotor RPM with hover tip Mach number of 0.628. A closed loop controller was designed to demonstrate trailing edge flap mean position and the peak to peak amplitude control. Further, a soft pitch link was designed and fabricated, to replace the stiff pitch link and thereby reduce the torsional stiffness of the blade to 2/rev. This soft pitch link allowed for blade root pitch motion in response to the trailing edge flap inputs. Blade pitch response due to both steady as well as sinusoidal flap deflections were demonstrated. Finally, tests were performed in Glenn L. Martin wind tunnel using a model rotor rig to assess the performance of motor-flap system in forward flight. A swashplateless trim using brushless DC motor actuated trailing edge flaps was achieved for a rotor operating at 1200 RPM and an advance ratio of 0.28. Also, preliminary exploration was carried out to test the scalability of the motor driven

Flyover noise is a problem that affects citizens, primarily those that live near or around places with high air traffic such as airports or military bases. Such noise can be of great annoyance. The focus of this thesis is in determining a method to create a high fidelity sound source simulation of rotorcraft noise for the purpose of producing a complete flyover scenario to be used in psychoacoustic testing. The focus of the sound source simulation is simulating rotorcraft noise fluctuations during level flight to aid in psychoacoustic testing to determine human perception of such noise. Current methods only model the stationary or time-average components when synthesizing the sound source. The synthesis process described in this thesis determines the steady-state waveform of the noise as well as the time-varying fluctuations for each rotor individually. The process explored in this thesis uses an empirical approach to synthesize flyover noise by directly using physical flyover recordings. Four different methods of synthesis were created to determine the combination of components that produce high fidelity sound source simulation. These four methods of synthesis are: a) Unmodulated main rotor b) Modulated main rotor c) Unmodulated main rotor combined with the unmodulated tail rotor d) Modulated main rotor combined with the modulated tail rotor Since the time-varying components of the source sound are important to the creation of high fidelity sound source simulation, five different types of time-varying fluctuations, or modulations, were implemented to determine the importance of the fluctuating components on the sound source simulation. The types of modulation investigated are a) no modulation, b) randomly applied generic modulation, c) coherently applied generic modulation, d) randomly applied specific modulation, and e) coherently applied specific modulation. Generic modulation is derived from a different section of the source recording to which it is applied. For

The idea of improved fight performance through changes in the control surfaces dates back to the advent of aviation with the Wright brothers' pioneering work on "wing warping," but it was not until the recent progress in material and actuator development that such control surfaces seemed practical for modern aircraft. This has opened the door to a new class of aircraft that have the ability to change shape or morph, which are being investigated due to the potential to have a single platform serve multiple mission objectives, as well as improve performance characteristics. While the majority of existing research for morphing aircraft has focused on fixedwing aircraft, rotary-wing aircraft have begun to receive more attention. The purpose of this body of work is to investigate the current state of morphing actuation technology for rotorcraft and improve upon it. Specifically, this work looks at two types of morphing: Pneumatic Artificial Muscle (PAM) actuated trailing edge flaps and conformal variable diameter morphing. First, active camber changes through the use of PAM powered trailing edge flaps were investigated due to the potential for reductions in power requirements and vibration/noise levels. A PAM based antagonistic actuation system was developed utilizing a novel combination of mechanism geometry and PAM bias contraction optimization to overcome the natural extension stiffening characteristics of PAMs. In open-loop bench-top testing against a "worst-case" constant torsional loading, the system demonstrated actuation authority suitable for both primary control and vibration/noise reduction. Additionally, closed-loop test data indicated that the system was capable of tracking complex waveforms consistent with those needed for rotorcraft control. This system demonstrated performance on-par with the state of the art pneumatic trailing edge flap actuators, yet with a much smaller footprint and impact on the rotor-blade. The second morphing system developed in thi

Time-resolved particle image and particle tracking velocimetry measurements were made in the particle-laden turbulent flow environment below a rotor hovering over a mobile sediment bed. The results were also compared to the near-wall flow produced by a nominally equivalent two-dimensional wall jet. The objective of the work was to understand the fluid dynamic mechanisms of how the mean flow, stochastic turbulence, and concentrated vorticity produced by the rotor affected the mobilization and pickup of particles from the sediment bed. Another objective was to better understand the assumptions that would be required for the development of models that are more applicable to rotor-induced particle mobilization. It was shown that the mean flow in the boundary layer at the ground below the rotor was similar to that of a wall jet. However, the instantaneous flow field and turbulence characteristics between these two flows were significantly different. Mobilized particles of 45--63 micron diameter (with a particle Reynolds number of less than 30 and a Stokes number of about 60) were individually identified and tracked, with the objective of relating any changes in the temporal evolution of the vortical flow and turbulence characteristics of the carrier flow phase to its coupling to the dispersed particle phase. The processes of particle mobilization and pickup from the bed were found to correlate to the Reynolds stresses and discrete turbulence events, respectively. The mean flow and turbulence characteristics were modified by the presence of particles in the near-wall region, showing clear evidence of two-way coupling between the phases of the resulting two-phase flow. Specifically, it was shown that the uplifted particles altered the carrier flow near the sediment bed, leading to an earlier distortion of the flow induced by the blade tip vortices and to the accelerated diffusion of the vorticity that they contained. The uplifted particles were also seen to modify the over

High-speed flow visualization and high-resolution particle image velocimetry experiments were conducted on a two-bladed rotor that was operated in a hovering state, both out of ground effect (well away from the ground) and in ground effect at several rotor heights. Recent advances in flow diagnostic instrumentation allowed measurements of the rotor wake to be performed with unprecedented levels of resolution. In particular, the goal of the present work was to gain a better understanding of the fluid dynamics of the wake sheets (and the blade tip vortices) that were trailed from the blades. The present work examined the effects produced by two blade sets: 1. A baseline untwisted blade, 2. A twisted blade with $-17^{\circ}$ of linear twist, and has revealed fluid dynamic details of the wake sheet that were hitherto unknown. For the measurements made with the rotor operating out of ground effect, the blade sets were tested at two blade loading coefficients of 0.053 and 0.08, although only the higher loading condition was tested with the rotors operating in ground effect. For the rotor operating out of ground effect, a helicoidal tip vortex was shown to form at the blade tip and the associated wake sheets were initially laid down as small-scale counterrotating vortical pairs. However, this initial vorticity quickly diffused and the sheet was then convected as a concentrated bands of turbulence, including several dominant eddies. Several types of sheet dynamics were documented in the rotor wake, including sheet interactions with the tip vortices, the detailed behavior of this interaction depending on both the blade twist and the rotor thrust. At earlier wake ages, a sinusoidal wave-like perturbation was seen to be formed on the wake sheets, although the growth in wave amplitude was limited as the sheets were convected and stretched in the velocity gradients in the downstream wake. When the rotor was operated in ground effect, the vorticity in the wake sheets persisted to

Brownout dust cloud simulations were conducted for rotorcraft undergoing representative landing maneuvers, primarily to examine the effects of different rotor placement and rotor/airframe configurations. The flow field generated by a helicopter rotor in ground effect operations was modeled by using an inviscid, incompressible, time-accurate Lagrangian free-vortex method, coupled to a semi-empirical approximation for the boundary layer flow near the ground. A surface singularity method was employed to represent the aerodynamic influence of a fuselage. A rigorous coupling strategy for the free-vortex method was developed to include the effects of rotors operating at different rotational speeds, such as a tail rotor. For the dispersed phase of the flow, particle tracking was used to model the dust cloud based on solutions to a decoupled form of the Basset-Boussinesq-Oseen equations appropriate to dilute gas particle suspensions of low Reynolds number Stokes flow. Important aspects of particle mobility and uplift in such vortically driven dust flows were modeled, which included a threshold-based model for sediment mobility and bombardment effects when previously suspended particles impact the bed and eject new particles. Various techniques were employed to reduce the computational cost of the dust cloud simulations, such as particle clustering and parallel programming using graphics processing units. The predicted flow fields near the ground and resulting dust clouds during the landing maneuvers were analyzed to better understand the physics behind their development, and to examine differences produced by various rotor and airframe configurations. Metrics based on particle counts and particle velocities in the field of view were developed to help quantify the severity of the computed brownout dust clouds. The presence of both a tail rotor and the fuselage was shown to cause both local and global changes to the aerodynamic environment near the ground and also influenced

To better understand the problem of rotor-induced particle motion and rotorcraft brownout, time-resolved, dual-phase particle image velocimetry and particle tracking velocimetry measurements were made in the flow produced by a small laboratory rotor that was hovering over a ground plane covered with a mobile sediment bed. To investigate the three-dimensionality of the wake and resultant particle field, flow measurements were made in vertical and horizontal planes around the rotor and near the ground. The primary goals of the work were to: 1. Characterize the fundamental flow physics of a rotor wake interacting with a sediment bed; 2. Investigate how rotor operating parameters, such as the disk loading, blade loading coefficient, and wake shedding frequency affected the mobilization, uplift and overall development of the particle field; 3. Examine the effects of placing a body between the rotor and the ground to understand how the interactions of the rotor wake with the body affected the transport of particles from the bed. The results showed that the rotor wake was very three-dimensional, with highly non-uniform velocities near the ground that resulted in the radially asymmetric mobilization, uplift and suspension of particles. The tip vortices were found to be the primary contributor to the uplift of particles, with the aperiodic variations in their trajectories near the ground causing intermittent particle mobilization events. These effects were caused, in part, by wave-like displacements that developed along the lengths of the tip vortices, which caused some parts of the filaments to convect closer to the ground than other parts and so uplift discrete bursts or plumes of particles. The quantity and distribution of uplifted particles were shown to be affected by the operating condition of the rotor, with the overall complexity of the rotor wake generally resulting in the formation of a highly three-dimensional and time-varying particle field. The rotor operating p

The action of the rotor wake on loose sediment on the ground is primarily responsible for inducing the rotorcraft brownout phenomenon. Therefore, any simulation of brownout must be capable of accurately predicting the velocity field induced by the rotor when it is operating in ground effect. This work attempts to use a compressible, structured, overset Reynolds-Averaged Navier-Stokes (RANS) based solver to simulate hovering rotors in ground effect (IGE) to demonstrate the capability of the code to provide accurate tip vortex flow field predictions, and provide a good understanding of the ground-wake interactions. The computations are performed for a micro-scale rotor (0.086m radius, aspect ratio of 4.387 operating at a tip Mach number of 0.08 and Reynolds number of 32,500) and a sub-scale rotor (0.408m radius, aspect ratio of 9.132 operating at a tip Mach number of 0.24 and Reynolds number of 250,000) in order to compare to experimental measurements. The micro-scale rotor has a rectangular tip shape and is simulated three rotor heights: 1.5R, 1.0R and 0.5R above ground (R = Rotor radius). The sub-scale rotor is simulated at one particular rotor height (i.e. 1R) but with four different tip shapes: rectangular, swept, BERP-like and slotted tip. Various mesh placement strategies are devised to efficiently capture the path of the tip vortices for both regimes. The micro-scale rotor simulations are performed using the Spalart Allmaras (S-A) turbulence model. The examination of the IGE tip vortex flow field suggests high degree of instabilities close to the ground. In addition, the induced velocities arising from the proximity of the rotor tip vortices causes flow separation at the ground. The sub-scale rotor simulations show a smeared out flow field even at early wake ages due to excessive turbulence levels. The distance function in the S-A turbulence model is modified using the Delayed Detached Eddy Simulation (DDES) approach and a correction to length scaling is includ

Conventional turbulence models calibrated for fully turbulent boundary layers often over-predict drag and heat transfer on aerodynamic surfaces with partially laminar boundary layers. A robust correlation-based model is developed for use in Reynolds-Averaged Navier-Stokes simulations to predict laminar-to-turbulent transition onset of boundary layers on external aerodynamic surfaces. The new model is derived from an existing transition model for the two-equation k-omega Shear Stress Transport (SST) turbulence model, and is coupled with the one-equation Spalart-Allmaras (SA) turbulence model. The transition model solves two transport equations for intermittency and transition momentum thickness Reynolds number. Experimental correlations and local mean flow quantities are used in the model to account for effects of freestream turbulence level and pressure gradients on transition onset location. Transition onset is triggered by activating intermittency production using a vorticity Reynolds number criterion. In the new model, production and destruction terms of the intermittency equation are modified to improve consistency in the fully turbulent boundary layer post-transition onset, as well as ensure insensitivity to freestream eddy viscosity value specified in the SA model. In the original model, intermittency was used to control production and destruction of turbulent kinetic energy. Whereas, in the new model, only the production of eddy viscosity in SA model is controlled, and the destruction term is not altered. Unlike the original model, the new model does not use an additional correction to intermittency for separation-induced transition. Accuracy of drag predictions are improved significantly with the use of the transition model for several two-dimensional single- and multi-element airfoil cases over a wide range of Reynolds numbers. The new model is able to predict the formation of stable and long laminar separation bubbles on low-Reynolds number airfoils that i