Aerospace Engineering Theses and Dissertations

Permanent URI for this collectionhttp://hdl.handle.net/1903/2737

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An Acoustic Analysis of Valveless Pulsejet Engines
(2015) Maqbool, Daanish; Cadou, Christopher P; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
Valveless pulsejets are extremely simple aircraft engines; essentially cleverly designed tubes with no moving parts. These engines utilize pressure waves, instead of machinery, for thrust generation, and have demonstrated thrust-to-weight ratios over 8 and thrust specific fuel consumption levels below 1 lbm/lbf-hr – performance levels that can rival many gas turbines. Despite their simplicity and competitive performance, they have not seen widespread application due to extremely high noise and vibration levels, which have persisted as an unresolved challenge primarily due to a lack of fundamental insight into the operation of these engines. This thesis develops two theories for pulsejet operation (both based on electro-acoustic analogies) that predict measurements better than any previous theory reported in the literature, and then uses them to devise and experimentally validate effective noise reduction strategies. The first theory analyzes valveless pulsejets as acoustic ducts with axially varying area and temperature. An electro-acoustic analogy is used to calculate longitudinal mode frequencies and shapes for prescribed area and temperature distributions inside an engine. Predicted operating frequencies match experimental values to within 6% with the use of appropriate end corrections. Mode shapes are predicted and used to develop strategies for suppressing higher modes that are responsible for much of the perceived noise. These strategies are verified experimentally and via comparison to existing models/data for valveless pulsejets in the literature. The second theory analyzes valveless pulsejets as acoustic systems/circuits in which each engine component is represented by an acoustic impedance. These are assembled to form an equivalent circuit for the engine that is solved to find the frequency response. The theory is used to predict the behavior of two interacting pulsejet engines. It is validated via comparison to experiment and data in the literature. The technique is then used to develop and experimentally verify a method for operating two engines in anti-phase without interfering with thrust production. Finally, Helmholtz resonators are used to suppress higher order modes that inhibit noise suppression via anti-phasing. Experiments show that the acoustic output of two resonator-equipped pulsejets operating in anti-phase is 9 dBA less than the acoustic output of a single pulsejet.
ACTIVE JET ACOUSTIC CONTROL OF LOW FREQUENCY, IN-PLANE HELICOPTER HARMONIC NOISE
(2012) Sargent, Daniel Caleb; Schmitz, Fredric H.; Baeder, James D.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
A new approach to reducing low frequency, in-plane harmonic noise of helicopter rotors is explored theoretically and experimentally in this dissertation. The active jet acoustic control methodology employs on-blade, tip located unsteady air blowing to produce an acoustic anti-noise waveform that reduces or cancels the observed noise at targeted positions in the acoustic far-field of the rotor system. This effectively reduces the distance at which the helicopter rotor can be aurally detected. An extended theoretical model of the subsonic air jet, which is modeled as both a source of mass and momentum, is presented. The model is applied to a baseline, full-scale, medium weight helicopter rotor for both steady and unsteady blowing. Significant reductions in low frequency, in-plane harmonic noise are shown to be possible for the theoretical rotor system by using physically reasonable unsteady jet velocities. A new model-scale active jet acoustic control experimental test rotor system is described in detail. Experimental measurements conducted in the University of Maryland Acoustic Chamber for the ~1/7th rotor, operated at a full-scale hover tip Mach number of 0.661, indicate that active jet acoustic control is a viable option for reducing low frequency, in-plane harmonic noise. Good correlation between theoretical predictions and measured data for four valve control cases are observed in both the time and frequency domains. Model-scale limitations of the tip-jet blowing experiment limited the peak noise level reductions to 30%. However, theory suggests that if the limitations of the model-scale controller are mitigated, much larger noise reductions are possible.
Aeroacoustic Analysis of Asymmetric Lift-Offset Helicopter in Forward Flight
(2021) Arias, Paulo; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
In recent years, the University of Maryland has worked on an asymmetric lift-offset compound helicopter. The configuration consists of a single main rotor helicopter with the addition of two key ways to increase the forward speed: a stubbed wing on the retreating fuselage side, and a slowed down rotor. Experiments and simulations have shown that the novel concept provides improved thrust potential and lift-to-drag ratios in high-speed forward flight. This study aims to determine whether the asymmetric lift-offset configuration also provides aeroacoustic benefits in forward flight in addition to its aerodynamic advantages. The aerodynamic results from previous computational and experimental studies are recreated using the Mercury framework, in-house Computational Fluid Dynamics solver based on Reynolds-Averaged Navier-Stokes (RANS) coupled to a comprehensive rotor analysis for structural deformations and trim. The acoustic analysis is performed using an acoustic code based on the Ffowcs William-Hawkings equation to solve for the tonal noise propagating from the surfaces of the aircraft. The BPM model is used for broadband noise prediction. It was found that for an advance ratio of 0.5 the wing-lift offset configuration can produce 56.8% more thrust at the same collective angle without any penalties in total noise. When the configurations produce equal thrust it was found that the wing-lift offset case has a 4 dB reduction in maximum overall sound pressure level. At an advance ratio of 0.3 with trim for equivalentthrust between configurations, a 3 dB maximum OASPL reduction was obtained with the inclusion of the wing. The rotor of the wing-lift offset case was also slowed down while maintaining equal thrust to find a 6 dB reduction at an advance ratio of 0.55. Blade flap and lag bending moments near the root were also significantly reduced for the wing-lift offset configuration with equal thrust.
Aeroacoustic Implications of Installed Propeller Interactional Aerodynamics and Transient Propeller Motions
(2023) Jayasundara, Dilhara; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
The emergence of advanced air mobility and sustainable aviation concepts have revived the interest in propeller-driven aircraft. A number of electric vertical take-off and landing (eVTOL) aircraft have been developed to cater to the demands of urban air mobility (UAM) and significant advancements have been made in unmanned aerial vehicles (UAV) equipped with vertical take-off and landing capabilities. However, the community acceptance of these new aircraft configurations highly depends on having a low noise footprint as they will operate in dense urban environments. Propeller noise is considered the major source of noise in these aircraft with the introduction of electric propulsion and it can significantly increase with the effects of installation and transient propeller motions. This study aims to comprehend the complex aerodynamic interactions within such aircraft that result from propeller installation and contribute to the generation of high noise levels. To understand the physics of propeller installation, a wingtip-mounted propeller was analyzed at several angles of attack using computational fluid dynamics (CFD) based on Reynolds-averaged Navier-Stokes (RANS) equations and computational aeroacoustics based on the Ffowcs Williams - Hawkings equation. The aeroacoustic implications of the propeller axis inclination and the propeller-wing aerodynamic interaction were studied in-depth. The propeller-wing interaction leads to a significant increase in propeller noise (~20 to 30 dB increase along the rotational axis) and causes the wing to generate a loading noise in the same order of magnitude as the propeller noise. To extrapolate the understanding of installation effects to a full aircraft, the aeroacoustic characteristics of a quadrotor biplane tailsitter were analyzed in both hover and forward flight focusing on the rotor-rotor and rotor-airframe aerodynamic interaction. The rotor-rotor interaction was found to be a significant source of loading noise in hover but having the fuselage as a physical barrier between the rotors largely reduces its effect. The airframe loading noise and rotor broadband noise are equally dominant as the rotor tonal noise when the aircraft is in forward flight. Moreover, the study evaluated the effectiveness of rotor synchrophasing in reducing the aircraft noise footprint and it showed promising results in hover, causing a reduction of aircraft noise by more than 10 dB. Furthermore, an efficient computational aeroacoustics framework was developed to facilitate the computations, ensuring optimal utilization of the computational resources. The CPU and GPU parallelization and other optimization techniques were able to achieve a 98% reduction in computation time for an isolated propeller case. This enabled the rapid aeroacoustic computations of periodic and non-periodic problems. This was used to analyze the aeroacoustics of an isolated propeller undergoing a transition from hover to forward flight. The aerodynamic and acoustic results of the unsteady case were compared with quasi-steady cases performed at intermediate tilt angles. The quasi-steady CFD simulations predicted the unsteady transition aeroacoustics with reasonable accuracy. A tilting quasi-steady approach was proposed to better capture the aerodynamics and acoustics of the unsteady transition.
Effect of Interactional Aerodynamics on Computational Aeroacoustics of Sikorsky's Notional X2 Platform
(2020) Bahr, Ian; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
An in-house acoustics code, ACUM, was used in conjunction with full vehicle CFD/CSD coupling to create a computational aeroacoustic framework to investigate the effect of aerodynamic interactions on the acoustic prediction of a compound coaxial helicopter. The full vehicle CFD/CSD was accomplished by using a high- fidelity computational fluid dynamics framework, HPCMP CREATETM-AV Helios, combined with an in-house computational structural dynamics solver to simulate the helicopter in steady forward flight. A notional X2TD helicopter consisting of a coaxial rotor, airframe, and pusher propeller was used and split into three simulation cases: isolated coaxial and propeller, airframe and full helicopter configuration to investigate each component’s effect on the others noise as well as the total noise. The primary impact on the acoustic prediction was the inclusion of the airframe in the CFD simulation as it affected both coaxial rotors as well as the propeller. It was found that the propeller and coaxial rotors had a negligible impact on each other.
Fundamental Rotorcraft Acoustic Modeling from Experiments (FRAME)
(2011) Greenwood, Eric; Schmitz, Fredric H; Hubbard, James E; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
A new methodology is developed for the construction of helicopter source noise models for use in mission planning tools from experimental measurements of helicopter external noise radiation. The models are constructed by employing a parameter identification method to an assumed analytical model of the rotor harmonic noise sources. This new method allows for the identification of individual rotor harmonic noise sources and allows them to be characterized in terms of their individual non-dimensional governing parameters. The method is applied to both wind tunnel measurements and ground noise measurements of two-bladed rotors. The method is shown to match the parametric trends of main rotor harmonic noise, allowing accurate estimates of the dominant rotorcraft noise sources to be made for operating conditions based on a small number of measurements taken at different operating conditions. The ability of this method to estimate changes in noise radiation due to changes in ambient conditions is also demonstrated.
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.
Quasi-Static Acoustic Mapping of Helicopter Blade-Vortex Interaction Noise
(2004-07-27) Gopalan, Gaurav; Schmitz, Fredric H; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
This research extends the applicability of storage-based noise prediction techniques to slowly maneuvering flight. The quasi-static equivalence between longitudinal decelerating flight and steady-state longitudinal descent flight, and its application to the estimation of BVI noise radiation under slow longitudinal maneuvering flight conditions, is investigated through various orders of flight dynamics modeling. The entire operating state of the helicopter is shown to be similar during equivalent flight conditions at the same flight velocity. This equivalence is also applied to the prediction of control requirements during longitudinal maneuvers. Inverse simulation based flight dynamics models of lower order are seen to capture many important trends associated with slow maneuvers, when compared with higher order modeling. The lower order flight dynamics model is used to design controlled maneuvers that may be practically flown during descent operations or as part of research flight testing. A version of a storage-based acoustic mapping technique, extended to slowly maneuvering longitudinal flight, is implemented for helicopter main rotor Blade-Vortex Interaction (BVI) noise. Various approach trajectories are formulated and analytical estimates of the BVI noise radiation characteristics associated with a full-scale two-bladed rotor are mapped to the ground using this quasi-static mapping approach. Multi-segment decelerating descent approaches are shown to be effective in ground noise abatement. The effects of steady longitudinal winds are investigated on radiated and ground noise. Piloting trim choices are seen to dominate the noise radiation under these flight conditions.
SURROGATE MODELING AND CHARACTERIZATION OF BLADE-WAKE INTERACTION NOISE FOR HOVERING SUAS ROTORS USING ARTIFICIAL NEURAL NETWORKS
(2022) Thurman, Christopher; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
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 subsequent rotor blade showed to be the most prominent type of blade-wake interaction noise, exhibiting an acoustic contribution upward of 7 dB. Blade-wake interaction noise caused by a direct impingement of a tip vortex on the leading edge of a subsequent rotor blade had the second largest acoustic significance, exhibiting roughly 6 dB of broadband noise. The third, and least significant type of blade-wake interaction noise was shown to be caused by impingement of a blade-wake sheet on the mid-span of a subsequent rotor. This last type of blade-wake interaction noise was seen to only occur in the turbulent-wake operating state and possibly mild vertical descent conditions, and had approximately a 2.5 dB acoustic contribution.