Helicopters are highly dependent on their transmission systems, which provide the vital links from the engines to the rotor and ancillary systems. Components are highly loaded and must be manufactured to a high degree of accuracy; the lack of redundancy implies that this is a 'series-chain' system. Existing techniques for calculating expected life are based upon historical data from different gearbox and helicopter types, thus limiting the confidence of the results. Design techniques may be conservative in some areas, whilst neglecting to consider different load patterns, usage, maintenance and environmental factors. This work describes the development of probabilistic models that represent damage accumulated by fatigue, wear and corrosion of the key components with an Intermediate gearbox (IGB). The parameters of these models represent geometrical, load and material data at the design stage, and produce an output in terms of failure probability against operating hours. This allows the influential parameters to be identified before building a prototype helicopter gearbox. The results from these models are then used to predict the upper and lower bounds of system reliability. This enables the combination of diverse failure mechanisms to be viewed to determine the relevant significance of each failure mechanisms. The effectiveness of the gearbox monitoring systems has been incorporated in the computer model by considering the probability of detection (POD) of each failure mechanism. The work to develop models found that there is a large body of work available to describe damage accumulation due to fatigue, but far less in regard to wear and corrosion. Fatigue models are very sensitive to load and material variability, particularly tooth root bending fatigue, for which many loads are considered 'non-damaging'. Wear models are mostly affected by changes in material hardness, wear coefficient and slip amplitude; changes in load are less influential on the predicted time to failure. The results for galvanic corrosion are dominated by the corrosion rate and time to initiate. In the system reliability model, reducing gear load appears to be the simplest means to increase life; increases in material strength and reduction in material variability are less achievable without significant improvements in manufacture and/or material technology.

Recent technological advances have lead to the development and application of multirotors for commercial package delivery and air taxis. They differ from helicopters because they operate at lower Reynolds numbers, induce more rotor-rotor interactions, and are controlled using variable-speed rather than variable-pitch rotors. The dynamic flow field of a multirotor leads to complex aerodynamics and aeroacoustics that can only be fully captured computationally using high-fidelity methods. A very popular high-fidelity approach used in the traditional rotorcraft simulation field is unsteady Reynolds-averaged Navier-Stokes (URANS). The URANS equations are implemented in this research using the CREATE-AV Helios computational code suite; however, when applying this method to smaller multirotors, three major challenges arise. The first challenge is to understand how the computational methods perform at the lower operational Reynolds numbers relevant to smaller multirotors. Implementation of URANS requires turbulence modeling, and models should account for laminar-turbulent transition at lower Reynolds numbers. This challenge is addressed by evaluating the effect of laminar-turbulent transition models in the Helios suite. This thesis demonstrates that current laminar-turbulent transition models in Helios do not improve either performance or noise predictions for small rotors. Fully-turbulent models provide reasonable accuracy at a lower cost. The second challenge to the high-fidelity multirotor simulation framework is the lack of an appropriate trim algorithm. For a computational prediction to be useful, it is necessary to simulate the rotorcraft in realistic flight conditions. This is accomplished via computation of trim, the set of controls and vehicle state that achieves a desired flight condition. Trim algorithms exist for helicopters that utilize blade pitch control but there has been no validation of high-fidelity computations that include trim for multirotors that control using individual rotor-speeds. This thesis presents the development and validation of a high-fidelity approach for multirotor trim using loose aerodynamic coupling. The newly developed trim algorithm currently solves unconstrained multirotor systems, which are underdetermined because the number of controls is greater than the number of targets, by adding an additional optimization for minimum power. The importance of including trim and performing the aerodynamic calculations with high-fidelity is demonstrated. The third challenge arises when applying high-fidelity methods to predict multirotor aeroacoustics. URANS alone cannot capture the full acoustic spectrum directly because the turbulent fluctuations are modeled and the grid cannot extend to the far field reliably. Therefore, it is necessary to develop models for portions of the spectrum and to use an acoustic analogy that converts sources in the CFD as noise to the far-field. However, there are complex interactional effects, a multitude of types of noise sources, and different methods for utilizing an acoustic analogy. It is also difficult to develop computations and experiments in which the noise sources match. The PSU-WOPWOP code is used to implement the Ffowcs-Williams and Hawkings equation with both impermeable and permeable surface approaches. Additionally, UCD-Quietfly is used to model some broadband noise sources. Through adaptation of existing tools and development of novel computational methods, the ability to predict multirotor noise is explored via computational investigation with comparison to experimental measurements to the extent possible. This thesis offers a new high-fidelity capability for predicting multirotor aerodynamics and aeroacoustics in trimmed flight. The approach enables final design analysis for developers of emerging multirotor systems who need reliable performance and noise predictions before full prototype development and testing. Capabilities developed in this thesis provide a platform upon which new trim optimization approaches can be created and tested, such as one that minimizes noise. / 2023-05-23T00:00:00Z

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 purposes of this study, it is not clearly dominated by either thickness or loading noise characteristics, but still displays long-term modulation. Random application of the modulation implies that there is a loss of absolute modulation phase and amplitude information across the frequency spectrum. Coherent application of the modulation implies that an attempt is made to line up the absolute phase and amplitude of the modulation signal with that which is being replaced (i.e. that which was stripped from the original recording and expanding or contracting to fit the signal to which it is applied). Specific modulation is the modulation from the source recording which is being reconstructed. A psychoacoustic test was performed to rank the fidelity of each synthesis method and each type of modulation. Performing this comparison for two different emission angles provides insight as to whether the ranking will differ between the emission angles. The modulated main rotor combined with the modulated tail rotor showed the highest fidelity and had a much higher fidelity than any of the other synthesis methods. The psychoacoustic test proved that modulation is necessary to produce a high fidelity sound source simulation. However, the use of a generic modulation or a randomly applied specific modulation proved to be an inadequate substitute for the coherently applied specific modulation. The results from this research show that more research is necessary to properly simulate a full flyover scenario. Specifically, more data is needed in order to properly model the modulation for level flight. / Master of Science

The Virginia Polytechnic Institute and State University has recently expanded its unmanned and autonomous systems research to include aerial vehicles. In the summer of 2004, members of the Autonomous Aerial Vehicle Team at Virginia Tech successfully competed in the Student UAV competition and the International Aerial Robotics Competition. The AAVT entered a fixed-wing vehicle in the former and a rotary-wing craft in the latter competition. Commercial flight controllers were used in both competitions in order to familiarize team members with this technology. The next step in research at VT focuses on developing an experimental rotorcraft platform to be used for control algorithm testing and development. Before the development of a flight control system is possible, a physical plant or model accurately describing the dynamics of the system is required. Use of the model in a virtual simulation environment is also beneficial to tune control gains and analyze robustness of the closed-loop system. The work presented focuses on developing a 6 degree-of-freedom model of a small-scale single shaft rotorcraft. The particular platform being developed is the Bergen Industrial Twin. In addition to dynamical concerns, attention is paid to performance characteristics of the aircraft. The nonlinear system of equations is solved which can be utilized in a simulated environment. Linear models are extracted and their control and stability characteristics are analyzed. Finally, the methodology is explained for obtaining models through system identification techniques using the CIFER facility. / Master of Science

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. / Master of Science / Helicopter 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 sounds heard by an observer as a helicopter flies overhead. 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. The current methods used to produce helicopter flyovers are limited to a predefined length of data in order to create sounds an individual may hear on the ground. It has been documented that creating flyover noise, from direct use of physical flyover recordings, yields a high fidelity signal, as long as all components are present when recreating the new sound. In order to extend these signals indefinitely, models must be developed in order to model the key components heard as a helicopter passes over an observer. The results from this work have successfully found a means in which to produce a viable model for certain components of the original flyover. Further investigation is still required in order to produce a high-fidelity model that can be extended indefinitely with all the necessary components included. This research is part of a broader effort to study the effects flyovers have on the population in terms of annoyance and detection. The work done here will help to aid further models used to determine what individuals find annoying with regard to helicopters and the noises they produce.

New technologies and new rotorcraft operations are being developed in order to meet new environmental requirements such as noise reduction and less pollutant emissions. In this project a parametric study was developed over a tilt rotor model in order to assess the environmental impact in terms of operational parameter and fuel burned looking at pollutant emission released into the air such as NOx, CO, UHC, PM, CO2 & H2O In order to perform the study previously stated, a computational tool build on Simulink titled tilt rotor mission performance was developed to run a single mission profile as a base line making different operational variations on every mission segment looking at deviations over fuel burned and pollutant emissions. The contribution of pollutant emissions during the cruise segment was compared to other phases obtaining 80% of CO2 and H2O, 75% of CO and UHC, 77% of NOx, and 78% of PM. Also, comparing the distance flown of the tilt rotor with some turboprop aircraft, it was found that the fuel burned and levels of CO2 are higher using tilt rotor rather than turboprop aircraft. On the other hand this is much better than helicopters.

The tumbling tip vortex effect of a reduced-scale, 1 m diameter, four-bladed rotor during hover is studied using vortex methods, combined with a center of mass analysis approach. Measurements of all three components of the velocity field are acquired using a stereo PIV system synchronized to capture up to 500 degrees of vortex age, with 10 degree wake age offsets, during hover conditions. The nominal operating condition of the rotor is at a rotational rate of 1520RPM, corresponding to ReC = 248,000 with a chord length of 58.5mm. The rotor is operated with a pitch of 7.2± 0.5 degrees and a CT/sigma of 0.045. The far wake vortex tumbling phenomenon is captured and described. It is shown that tip vortices from two blades tumble through approximately 90 degrees of rotation before they coalesce. It is also seen that the constituent parent vortices do not combine to create a stronger daughter vortex as was previously thought to happen. Instead, the merged vortex has a lower large-radius circulation than either of its parent vortices. An accurate characterization and prediction of the trajectory of the far wake vortex tumbling can enhance the ability to predict and alleviate the resuspension of particles during brownout as well as provide a database for far wake validation of CFD codes. / text

Modern helicopters, civilian and military alike, are expected to operate in all weather conditions. Ice accretion adversely affects the availability, affordability, safety and survivability. Availability of the vehicle may be compromised if the ice formation requires excessive torque to overcome the drag needed to operate the rotor. Affordability is affected by the power requirements and cost of ownership of the deicing systems needed to safely operate the vehicle. Equipment of the rotor blades with built-in heaters greatly increases the cost of the helicopter and places further demands on the engine. The safety of the vehicle is also compromised due to ice shedding events, and the onset of abrupt, unexpected stall phenomena attributable to ice formation. Given the importance of understanding the effects of icing on aircraft performance and certification, considerable work has been done on the development of analytical and empirical tools, accompanied by high quality wind tunnel and flight test data. In this work, numerical studies to improve ice growth modeling have been done by reducing limitations and empiricism inherent in existing ice accretion models. In order to overcome the weakness of Lagrangian approach in unsteady problem such as rotating blades, a water droplet solver based on 3-D Eulerian method is developed and integrated into existing CFD solver. Also, the differences between the industry standard ice accretion analyses such as LEWICE and the ice accretion models based on the extended Messinger model are investigated through a number of 2-D airfoil and 3-D rotor blade ice accretion studies. The developed ice accretion module based on 3-D Eulerian water droplet method and the extended Messinger model is also coupled with an existing empirical ice shedding model. A series of progressively challenging simulations have been carried out. These include ability of the solvers to model airloads over an airfoil with a prescribed/simulated ice shape, collection efficiency modeling, ice growth, ice shedding, de-icing modeling, and assessment of the degradation of airfoil or rotor performance associated with the ice formation. While these numerical simulation results are encouraging, much additional work remains in modeling detailed physics important to rotorcraft icing phenomena. Despite these difficulties, progress in assessing helicopter ice accretion has been made and tools for initial analyses have been developed.Modern helicopters, civilian and military alike, are expected to operate in all weather conditions. Ice accretion adversely affects the availability, affordability, safety and survivability. Availability of the vehicle may be compromised if the ice formation requires excessive torque to overcome the drag needed to operate the rotor. Affordability is affected by the power requirements and cost of ownership of the deicing systems needed to safely operate the vehicle. Equipment of the rotor blades with built-in heaters greatly increases the cost of the helicopter and places further demands on the engine. The safety of the vehicle is also compromised due to ice shedding events, and the onset of abrupt, unexpected stall phenomena attributable to ice formation. Given the importance of understanding the effects of icing on aircraft performance and certification, considerable work has been done on the development of analytical and empirical tools, accompanied by high quality wind tunnel and flight test data. In this work, numerical studies to improve ice growth modeling have been done by reducing limitations and empiricism inherent in existing ice accretion models. In order to overcome the weakness of Lagrangian approach in unsteady problem such as rotating blades, a water droplet solver based on 3-D Eulerian method is developed and integrated into existing CFD solver. Also, the differences between the industry standard ice accretion analyses such as LEWICE and the ice accretion models based on the extended Messinger model are investigated through a number of 2-D airfoil and 3-D rotor blade ice accretion studies. The developed ice accretion module based on 3-D Eulerian water droplet method and the extended Messinger model is also coupled with an existing empirical ice shedding model. A series of progressively challenging simulations have been carried out. These include ability of the solvers to model airloads over an airfoil with a prescribed/simulated ice shape, collection efficiency modeling, ice growth, ice shedding, de-icing modeling, and assessment of the degradation of airfoil or rotor performance associated with the ice formation. While these numerical simulation results are encouraging, much additional work remains in modeling detailed physics important to rotorcraft icing phenomena. Despite these difficulties, progress in assessing helicopter ice accretion has been made and tools for initial analyses have been developed.

A new unsteady rotor/fuselage interactional aerodynamics model has been developed. This model loosely couples a Generalized Dynamic Wake Theory (GDWT) to a Navier-Stokes solution procedure. This coupling is achieved using a newly developed unsteady pressure jump boundary condition in the Navier-Stokes model. The new unsteady pressure jump boundary condition models each rotor blade as a moving pressure jump which travels around the rotor azimuth =and is applied between two adjacent planes in a cylindrical, non-rotating grid. Comparisons are made between predictions using this new model and experiments for an isolated rotor and for a coupled rotor/fuselage configuration. / Ph. D.

The Active Control of Structural Response (ACSR) vibration suppression system, where hydraulic actuators located between the gearbox and the fuselage are used to cancel vibration in large helicopters, has been used successfully for many years. However the power consumed by the actuators can be high, and using hydraulic actuation for smaller rotorcraft has not been seen as practical. In contrast to active vibration reduction systems, passive vibration isolation systems require no external power. Passive vibration isolation systems however have the disadvantage of being limited to working at one specific frequency which will not be acceptable as slowed rotor flight becomes more common for fuel efficiency and noise legislation reasons. In this thesis two electrically powered actuation concepts, one piezoelectric, and one electromagnetic were initially evaluated. An electrically powered actively augmented passive, or hybrid, vibration reduction system based on an electro hydrostatic actuator (EHA) concept was proposed to be developed further. This hybrid actuator will have a wider range of operating frequencies than a purely passive system, and have lower power consumption than a purely active system. The design is termed a “Resonant EHA”; in that the resonant frequency of the coupled fluid, pump and electric motor rotor inertia matches the fundamental vibration frequency. The hydraulic cylinder, fluid and pump act as a single stage gear ratio, and the. brushless electric motor’s inertia is the main resonating mass as in a Dynamic Antiresonant Vibration Isolator (DAVI) passive vibration reduction system. The electrical power is used to compensate for friction in the actuator and other losses, and if needed can shift the operating point away from the resonant frequency. Simulation results indicated that a hydraulic circuit in which the pump leakage is fed back into the low pressure line would introduce unacceptable disturbances in the flows to and from the cylinder. To eliminate the source of the disturbances, a fully integrated electric motor and pump circuit design was chosen in which the electric motor is immersed in hydraulic fluid. An EHA demonstrator was built sized for a 1.5 tonne rotorcraft. For sizing comparison purposes the frameless brushless D.C motor for each strut of 1.5 tonne rotorcraft has a rotor and stator mass of approximately 1 kg, and can produce a continuous stall torque of 2 Nm. The bidirectional pump has a displacement of 1.5 cm3/rev, the mean system pressure was taken as 90 bar, and the double ended hydraulic cylinder has a 32 mm diameter bore, and 18 mm rod. Initial test results for the proof of concept EHA showed highly significant free play with a reversal of torque direction, resulting in unacceptable loss in transmission stiffness. The free play was traced to the gear pump and a hypothesis for the origin of the free play was put forward. To avoid torque reversals the EHA was further tested with a constant offset torque bias which proved successful in restoring a sufficient stiffness to the transmission. The sizing of the electric motor and power consumed with a non-zero offset torque is greater than a torque reversing motor, which limits the immediate application of the device in the present form. Future research investigating the use of other transmission elements, such as a piston pump, to obtain a more linear stiffness is recommended. As a hybrid vibration isolation system a Root Mean Square (RMS) reduction by a factor of four and near elimination of the fundamental frequency vibrations was achieved for the frequency range of 10 to 20 hertz.