UMD Theses and Dissertations
Permanent URI for this collectionhttp://hdl.handle.net/1903/3
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Item Effect of Sloped Terrain on In-Ground-Effect Hover Performance for an Isolated Rotor(2023) Prewitt, Jack; Tritschler, John; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)The present work conducted performance testing using a laboratory-scale isolated rotor operating over a ground plane mounted to a six degree-of-freedom motion platform to simulate in-ground-effect operations over sloped terrain. The rotor utilized a pair of 1:13.24 scale OH-58C blades, and performance measurements were collected using a six-axis load cell to which the rotor was mounted. Seven ground plane angles ranging from 0–18 deg, five collective blade pitches ranging from 0–8 deg, and 15 hub heights ranging from a nondimensional hub height, z/R, of 0.6 to 2.0 were tested. Additionally, the rotor was operated out-of-ground-effect for collective blade pitches ranging from 0–12 deg in increments of 1 deg in order to compare in-ground-effect and out-of-ground-effect hover performance. In-ground-effect hover over sloped terrain was found to have over a 7% reduction in performance as compared to hover over level terrain at low hub heights and large ground plane angles. In-ground-effect hover over sloped terrain was also found to require 2% more power than hover out-of-ground-effect at high hub heights and large ground plane angles. Finally, a semi-empirical model for hover performanceover sloped terrain was developed on the basis of the classic Cheeseman and Bennett ground effect model for level terrain. The coefficients obtained from this model were found to change in a consistent manner as both ground plane angle and blade loading coefficient changed, which suggests that the model could be used for future performance predictions for hover over sloped terrain.Item NUMERICAL SIMULATION AND VALIDATION OF HELICOPTER BLADE-VORTEX INTERACTION USING COUPLED CFD/CSD AND THREE LEVELS OF AERODYNAMIC MODELING(2014) Amiraux, Mathieu; Baeder, James D; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Rotorcraft Blade-Vortex Interaction (BVI) remains one of the most challenging flow phenomenon to simulate numerically. Over the past decade, the HART-II rotor test and its extensive experimental dataset has been a major database for validation of CFD codes. Its strong BVI signature, with high levels of intrusive noise and vibrations, makes it a difficult test for computational methods. The main challenge is to accurately capture and preserve the vortices which interact with the rotor, while predicting correct blade deformations and loading. This doctoral dissertation presents the application of a coupled CFD/CSD methodology to the problem of helicopter BVI and compares three levels of fidelity for aerodynamic modeling: a hybrid lifting-line/free-wake (wake coupling) method, with modified compressible unsteady model; a hybrid URANS/free-wake method; and a URANS-based wake capturing method, using multiple overset meshes to capture the entire flow field. To further increase numerical correlation, three helicopter fuselage models are implemented in the framework. The first is a high resolution 3D GPU panel code; the second is an immersed boundary based method, with 3D elliptic grid adaption; the last one uses a body-fitted, curvilinear fuselage mesh. The main contribution of this work is the implementation and systematic comparison of multiple numerical methods to perform BVI modeling. The trade-offs between solution accuracy and computational cost are highlighted for the different approaches. Various improvements have been made to each code to enhance physical fidelity, while advanced technologies, such as GPU computing, have been employed to increase efficiency. The resulting numerical setup covers all aspects of the simulation creating a truly multi-fidelity and multi-physics framework. Overall, the wake capturing approach showed the best BVI phasing correlation and good blade deflection predictions, with slightly under-predicted aerodynamic loading magnitudes. However, it proved to be much more expensive than the other two methods. Wake coupling with RANS solver had very good loading magnitude predictions, and therefore good acoustic intensities, with acceptable computational cost. The lifting-line based technique often had over-predicted aerodynamic levels, due to the degree of empiricism of the model, but its very short run-times, thanks to GPU technology, makes it a very attractive approach.Item Modeling Helicopter Near-Horizon Harmonic Noise due to Transient Maneuvers(2013) Sickenberger, Richard Dwight; Baeder, James; Schmitz, Fredric; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)A new first principles model has been developed to estimate the external harmonic noise radiation for a helicopter performing transient maneuvers in the longitudinal plane. This model, which simulates the longitudinal fuselage dynamics, main rotor blade flapping, and far field acoustics, was validated using in-flight measurements and recordings from ground microphones during a full-scale flight test featuring a Bell 206B-3 helicopter. The flight test was specifically designed to study transient maneuvers. The validated model demonstrated that the flapping of the main rotor blades does not significantly affect the acoustics radiated by the helicopter during maneuvering flight. Furthermore, the model also demonstrated that Quasi-Static Acoustic Mapping (Q-SAM) methods can be used to reliably predict the noise radiated during transient maneuvers. The model was also used to identify and quantify the contributions of main rotor thickness noise, low frequency loading noise, and blade-vortex interaction (BVI) noise during maneuvering flight for the Bell 206B-3 helicopter. Pull-up and push-over maneuvers from pure longitudinal cyclic and pure collective control inputs were investigated. The contribution of thickness noise and low frequency loading noise during maneuvering flight was found to depend on the orientation of the tip-path plane relative to the observer. The contribution of impulsive BVI noise during maneuvering flight was found to depend on the inflow through the main rotor and the orientation of the tip-path plane relative to the observer.Item Design and Performance Prediction of Swashplateless Helicopter Rotors with Trailing Edge Flaps and Tabs(2010) Falls, Jaye; Chopra, Inderjit; Datta, Anubhav; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)This work studies the design of trailing edge controls for swashplateless helicopter primary control, and examines the impact of those controls on the performance of the rotor. The objective is to develop a comprehensive aeroelastic analysis for swashplateless rotors in steady level flight. The two key issues to be solved for this swashplateless control concept are actuation of the trailing edge controls and evaluating the performance of the swashplateless rotor compared to conventionally controlled helicopters. Solving the first requires simultaneous minimization of trailing flap control angles and hinge moments to reduce actuation power. The second issue requires not only the accurate assessment of swashplateless rotor power, but also similar or improved performance compared to conventional rotors. The analysis consists of two major parts, the structural model and the aerodynamic model. The inertial contributions of the trailing edge flap and tab are derived and added to the system equations in the structural model. Two different aerodynamic models are used in the analysis, a quasi-steady thin airfoil theory that includes arbitrary hinge positions for the flap and the tab, and an unsteady lifting line model with airfoil table lookup based on wind tunnel test data and computational fluid dynamics simulation. The predicted swashplateless rotor power is sensitive to the pattern of trailed vorticity from the rotor blade. Trailed vortices are added at the inboard and outboard boundaries of the trailing edge flap, and the flap deflection is used to calculate an effective angle of attack for the calculation of the near and far wake. This wake model predicts the swashplateless rotor requires less main rotor power than the conventional UH-60A helicopter from hover to &mu = 0.25. As the forward flight speed increases, the swashplateless predicted power increases above the conventional rotor, and the rotor lift-to-drag ratio decreases below that of the conventional rotor.Item Contributions to the dynamics of helicopters with active rotor controls(2008-07-15) Malpica, Carlos; Celi, Roberto; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)This dissertation presents an aeromechanical closed loop stability and response analysis of a hingeless rotor helicopter with a Higher Harmonic Control (HHC) system for vibration reduction. The analysis includes the rigid body dynamics of the helicopter and blade flexibility. The gain matrix is assumed to be fixed and computed off-line. The discrete elements of the HHC control loop are rigorously modeled, including the presence of two different time scales in the loop. By also formulating the coupled rotor-fuselage dynamics in discrete form, the entire coupled helicopter-HHC system could be rigorously modeled as a discrete system. The effect of the periodicity of the equations of motion is rigorously taken into account by converting the system into an equivalent system with constant coefficients and identical stability properties using a time lifting technique. The most important conclusion of the present study is that the discrete elements in the HHC loop must be modeled in any HHC analysis. Not doing so is unconservative. For the helicopter configuration and HHC structure used in this study, an approximate continuous modeling of the HHC system indicates that the closed loop, coupled helicopter-HHC system remains stable for optimal feedback control configurations which the more rigorous discrete analysis shows can result in closed loop instabilities. The HHC gains must be reduced to account for the loss of gain margin brought about by the discrete elements. Other conclusions of the study are: (i) the HHC is effective in quickly reducing vibrations, at least at its design condition, although the time constants associated with the closed loop transient response indicate closed loop bandwidth to be 1~rad/sec on average, thus overlapping with FCS or pilot bandwidths, and raising the issue of potential interactions; (ii) a linearized model of helicopter dynamics is adequate for HHC design, as long as the periodicity of the system is correctly taken into account, i.e., periodicity is more important than nonlinearity, at least for the mathematical model used in this study; and (iii) when discrete and continuous systems are both stable, and quasisteady conditions can be guaranteed, the predicted HHC control harmonics are in good agreement.Item Development and Investigation of a Flapping Rotor for Micro Air Vehicles(2007-07-09) Fitchett, Brandon Kurt; Chopra, Inderjit; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)This thesis describes the concept, design and testing of a micro air vehicle rotor testbed capable of independently controlled blade rotation and powered blade flapping. The design, dubbed the "Flotor", combined the benefits of a conventional MAV helicopter rotor with avian based flapping motion. The Flotor was tested as a conventional rotor, a conventional rotor with powered blade flapping, and a torqueless, freely rotating rotor with powered blade flapping. As a conventional rotor with a maximum figure of merit of 0.5, the results from the Flotor were similar to previously published experiments. With conventional rotation plus powered blade flapping at up to 8 per rotor revolution at a reduced frequency of 0.6, the maximum thrust increased by up to 15% due to delayed stall. The torque required at moderate thrust levels was reduced by up to 30%. The results from a 2-D quasi-steady blade element momentum analysis predicted average rotor loads accurately below 20° collective. As the first attempt at a torqueless flapping MAV rotor, the Flotor was capable of producing thrust and blade loadings comparable to flying animals, but less than current MAVs.Item Helicopter Flight Dynamics Simulation with a Time-Accurate Free-Vortex Wake Model(2007-04-26) Ribera, Maria; Celi, Roberto; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)This dissertation describes the implementation and validation of a coupled rotor-fuselage simulation model with a time-accurate free-vortex wake model capable of capturing the response to maneuvers of arbitrary amplitude. The resulting model has been used to analyze different flight conditions, including both steady and transient maneuvers. The flight dynamics model is based on a system of coupled nonlinear rotor-fuselage differential equations in first-order, state-space form. The rotor model includes flexible blades, with coupled flap-lag-torsion dynamics and swept tips; the rigid body dynamics are modeled with the non-linear Euler equations. The free wake models the rotor flow field by tracking the vortices released at the blade tips. Their behavior is described by the equations of vorticity transport, which is approximated using finite differences, and solved using a time-accurate numerical scheme. The flight dynamics model can be solved as a system of non-linear algebraic trim equations to determine the steady state solution, or integrated in time in response to pilot-applied controls. This study also implements new approaches to reduce the prohibitive computational costs associated with such complex models without losing accuracy. The mathematical model was validated for trim conditions in level flight, turns, climbs and descents. The results obtained correlate well with flight test data, both in level flight as well as turning and climbing and descending flight. The swept tip model was also found to improve the trim predictions, particularly at high speed. The behavior of the rigid body and the rotor blade dynamics were also studied and related to the aerodynamic load distributions obtained with the free wake induced velocities. The model was also validated in a lateral maneuver from hover. The results show improvements in the on-axis prediction, and indicate a possible relation between the off-axis prediction and the lack of rotor-body interaction aerodynamics. The swept blade model improves both the on-axis and off-axis response. An axial descent though the vortex ring state was simulated. As the"vortex ring" goes through the rotor, the unsteady loads produce large attitude changes, unsteady flapping, fluctuating thrust and an increase in power required. A roll reversal maneuver was found useful in understanding the cross-couplings effects found in rotorcraft, specifically the effect of the aerodynamic loading on the rotor orientation and the off-axis response.Item Development of Mach Scale Rotors with Composite Tailored Couplings for Vibration Reduction(2004-11-29) Bao, Jinsong; Chopra, Inderjit; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)The use of composite tailored couplings in rotor blades to reduce vibratory hub loads was studied through design, structural and aeroelastic analysis, fabrication, and wind tunnel test of Mach scale articulated composite rotors with tailored flap-bending/torsion couplings. The rotor design was nominally based on the UH-60 BLACK HAWK rotor. The 6-foot diameter blades have a SC1095 profile and feature a linear twist of -12 deg. The analysis of composite rotor was carried out using a mixed cross-section structural model, and UMARC. Five sets of composite rotor were fabricated, including a baseline rotor without coupling, rotors with spanwise uniform positive coupling and negative coupling, and rotors with spanwise dual-segmented coupling (FBT-P/N) and triple-segmented coupling. The blade composite D-spar is the primary structural element supporting the blade loads and providing the desired elastic couplings. Non-rotating tests were performed to examine blade structural properties. The measurements showed good correlation with predictions, and good repeatability for the four blades of each rotor set. All rotors were tested at a rotor speed of 2300 rpm (tip Mach number 0.65) at different advance ratios and thrust levels, in the Glenn L. Martin Wind Tunnel at the University of Maryland. The test results showed that flap-bending/torsion couplings have a significant effect on the rotor vibratory hub loads. All coupled rotors reduced the 4/rev vertical force for advance ratios up to 0.3, with reductions ranging from 1 to 34%. The mixed coupling rotor FBT-P/N reduced overall 4/rev hub loads at advance ratios of 0.1, 0.2 and 0.3. At a rotor speed of 2300 rpm and an advance ratio of 0.3, the FBT-P/N rotor achieved 15% reduction for 4/rev vertical force, 3% for 4/rev in-plane force and 14% for 4/rev head moment. The reductions in the 4/rev hub loads are related to the experimentally observed reductions in 3/rev and 5/rev blade flap bending moments. Through the present research, it has been experimentally demonstrated that structural couplings can significantly impact rotor vibration characteristics, and with suitable design optimization (coupling strength and spanwise distribution) they can be used to reduce vibratory hub loads without penalties.