Active and Passive Helicopter Noise Reduction Using the AVINOR/HELINOIR Code Suite

Peer Reviewedhttps://deepblue.lib.umich.edu/bitstream/2027.42/143085/1/1.C034519.pd

Helicopter tail rotor orthogonal blade vortex interaction

The aerodynamic operating environment of the helicopter is particularly complex and, to some extent, dominated by the vortices trailed from the main and tail rotors. These vortices not only determine the form of the induced flow field but also interact with each other and with elements of the physical structure of the flight vehicle. Such interactions can have implications in terms of structural vibration, noise generation and flight performance. In this paper, the interaction of main rotor vortices with the helicopter tail rotor is considered and, in particular, the limiting case of the orthogonal interaction. The significance of the topic is introduced by highlighting the operational issues for helicopters arising from tail rotor interactions. The basic phenomenon is then described before experimental studies of the interaction are presented. Progress in numerical modelling is then considered and, finally, the prospects for future research in the area are discussed

Langley aerospace test highlights, 1985

The role of the Langley Research Center is to perform basic and applied research necessary for the advancement of aeronautics and space flight, to generate new and advanced concepts for the accomplishment of related national goals, and to provide research advice, technological support, and assistance to other NASA installations, other government agencies, and industry. Significant tests which were performed during calendar year 1985 in Langley test facilities, are highlighted. Both the broad range of the research and technology activities at the Langley Research Center and the contributions of this work toward maintaining United States leadership in aeronautics and space research, are illustrated. Other highlights of Langley research and technology for 1985 are described in Research and Technology-1985 Annual Report of the Langley Research Center

Acoustics Division recent accomplishments and research plans

The research program currently being implemented by the Acoustics Division of NASA Langley Research Center is described. The scope, focus, and thrusts of the research are discussed and illustrated for each technical area by examples of recent technical accomplishments. Included is a list of publications for the last two calendar years. The organization, staff, and facilities are also briefly described

A multi-fidelity framework for physics based rotor blade simulation and optimization

New helicopter rotor designs are desired that offer increased efficiency, reduced vibration, and reduced noise. This problem is multidisciplinary, requiring knowledge of structural dynamics, aerodynamics, and aeroacoustics. Rotor optimization requires achieving multiple, often conflicting objectives. There is no longer a single optimum but rather an optimal trade-off space, the Pareto Frontier. Rotor Designers in industry need methods that allow the most accurate simulation tools available to search for Pareto designs. Computer simulation and optimization of rotors have been advanced by the development of "comprehensive" rotorcraft analysis tools. These tools perform aeroelastic analysis using Computational Structural Dynamics (CSD). Though useful in optimization, these tools lack built-in high fidelity aerodynamic models. The most accurate rotor simulations utilize Computational Fluid Dynamics (CFD) coupled to the CSD of a comprehensive code, but are generally considered too time consuming where numerous simulations are required like rotor optimization. An approach is needed where high fidelity CFD/CSD simulation can be routinely used in design optimization. This thesis documents the development of physics based rotor simulation frameworks. A low fidelity model uses a comprehensive code with simplified aerodynamics. A high fidelity model uses a parallel processor capable CFD/CSD methodology. Both frameworks include an aeroacoustic simulation for prediction of noise. A synergistic process is developed that uses both frameworks together to build approximate models of important high fidelity metrics as functions of certain design variables. To test this process, a 4-bladed hingeless rotor model is used as a baseline. The design variables investigated include tip geometry and spanwise twist. Approximation models are built for high fidelity metrics related to rotor efficiency and vibration. Optimization using the approximation models found the designs having maximum rotor efficiency and minimum vibration. Various Pareto generation methods are used to find frontier designs between these two anchor designs. The Pareto anchors are tested in the high fidelity simulation and shown to be good designs, providing evidence that the process has merit. Ultimately, this process can be utilized by industry rotor designers with their existing tools to bring high fidelity analysis into the preliminary design stage of rotors.Ph.D.Committee Co-Chair: Dr. Dimitri Mavris; Committee Co-Chair: Dr. Lakshmi N. Sankar; Committee Member: Dr. Daniel P. Schrage; Committee Member: Dr. Kenneth S. Brentner; Committee Member: Dr. Mark Costell

Rotorcraft In-Plane Noise Reduction Using Active/Passive Approaches With Induced Vibration Tracking

A comprehensive study of the use of active and passive approaches for in-plane noise reduction, including the vibrations induced during noise reduction, was conducted on a hingeless rotor configuration resembling the MBB BO-105 rotor. First, a parametric study was performed to examine the effects of rotor blade stiffness on the vibration and noise reduction performance of a 20%c plain trailing edge flap and a 1.5%c sliding microflap. This was accomplished using a comprehensive code AVINOR (for Active VIbration and NOise Reduction). A two-dimensional unsteady reduced order aerodynamic model (ROM), using the Rational Function Approximation approach and CFD-based oscillatory aerodynamic load data, was used in the comprehensive code. The study identified a hingeless blade configuration with torsional frequency of 3.17/rev as an optimum configuration for studying vibration and noise reduction using on-blade control devices such as flaps or microflaps. Subsequently, a new suite of computational tools capable of predicting in-plane low frequency sound pressure level (LFSPL) rotorcraft noise and its control was developed, replacing the acoustic module WOPWOP in AVINOR with a new acoustic module HELINOIR (for HELIcopter NOIse Reduction), which overcomes certain limitations associated with WOPWOP. The new suite, consisting of the AVINOR/HELINOIR combination, was used to study active flaps, as well as microflaps operating in closed-loop mode for in-plane noise reduction. An alternative passive in-plane noise reduction approach using modification to the blade tip in the 10%R outboard region was also studied. The new suite consisting of the AVINOR/HELINOIR combination based on a compact aeroacoustic model was validated by comparing with wind tunnel test results, and subsequently verified by comparing with computational results. For active control, the in-plane noise reduction obtained with a single 20%c plain trailing edge flap during level flight at a moderate advance ratio was examined. Different configurations of far-field and near-field feedback microphone locations were examined to develop a fundamental understanding of the feedback microphone locations on the noise reduction process A near-field microphone located on the tip of a nose boom was found to produce a LFSPL reduction of up to 6dB. However, this noise reduction was accompanied by an out-of-plane noise increase of 18dB and 60% increase in vertical hub shear. For passive control, three tip geometries having sweep, dihedral, and anhedral, were considered. The tip dihedral reduced LFSPL by up to 2dB without a vibratory load penalty. However, this was accompanied by an increase in the mid frequency sound pressure levels (MFSPL). The tip sweep and tip anhedral produced an increase in in-plane LFSPL below the horizon. A comparison of the active and passive approaches indicated that active approaches implemented by a plain flap with a feedback microphone located on the nose boom is superior to the passive control approaches. However, there is a general trade-off between LFSPL reduction, MFSPL generation and vibratory hub loads induced by noise control.PHDAerospace EngineeringUniversity of Michigan, Horace H. Rackham School of Graduate Studieshttps://deepblue.lib.umich.edu/bitstream/2027.42/138495/1/cmianghw_1.pd

Investigation of orthogonal blade-vortex interaction using a particle image velocimetry technique

The complex flowfield which is associated with a rotor wake gives rise to the multitude of aerodynamic interactions that may occur during rotorcraft operation. These interactions may give rise to undesirable noise and lead to an unacceptable performance degradation, and as such the investigation of the fundamental mechanics of such interactions, that which occurs between the tail rotor and the trailing tip vortices shed from the main rotor assembly, is the focus of the current investigation. As the purpose of the tail rotor is to provide balance for the torque of the main rotor, these types of interaction will adversely impact on the overall rotorcraft performance. The basis of the present thesis has been an experimental investigation of the orthogonal BVI, in which the axis of the interacting vortex (in the plane of the vortex core axial flow) is nominally orthogonal to the interacting blade chordline, representing the tail rotor interaction. The tests have been conducted using a specifically designed facility at the University of Glasgow, with the flow interrogated using a Particle Image Velocimetry (PVI) technique. The PVI method allows global flowfield information to be obtained pertaining to the nature of the interaction. The methodology was benchmarked against synthetic flowfields, and with the accuracy of the flowfield measurements improved dramatically with the implementation of the Forward/Reverse Tile Test (FRTT), which improved the accuracy in the flowfields to 3% in two-dimensional interrigation, and 5% in three-dimensional. The interrogation of the flowfield around the representative tail rotor blade demonstrated that the characteristics imparted vortex due to the BVI event could be attributed to the manner in which the axial flow component of the vortex was affected by the interaction. The results for the isolated flow conditions agreed well with those from previous measurements of the vortical structure, and the post interaction structure clearly indicated distinct differences determined by the direction of the axial flow relative to the blade chordline. Initial testing indicated that the thickness ratio had a marked effect on the progression of the OBVI, and for a suitably high thickness ratio, there was little evidence to suggest that the vortex core axial flow is 'cut' by the interacting body in the manner observed for the lower thickness ratios. For lower thickness ratios, as the vortex core is blocked by the interacting blade surface, the retardation of the axial component on the blade lower surface leads to rapid redistribution of the fluid into the surrounding flow, and the corresponding enlargement and distortion to the vortex tangential velocity components promoted by the radial outflow. On the upper blade side, regions of negative axial flow velocity indicate the presence of some fluid passing down through the core towards the surface of the blade, which are accompanied by a split divergence pattern around the vortex core. The effects immediately behind the trailing edge continue to be of interest due to the manner in which the vortex might be regenerated after the interaction and before any subsequent interactions with following blades. A relative lack of distortion within the out-of-plane component indicates that a rapid regeneration of the axial flow component may occur once the vortex has passed over the trailing edge. The use of passive control techniques in reduction of the effects associated with the orthoganal BVI have also been addressed, considering the effect of a counter-rotating vortex pair on the progression of the interaction. Although the inclusion a notch in the leading edge and outboard sweep on the rotor blade producing the representative trailing tip vortex did produce a well defined inboard vortex structure, there is evidence to suggest that this structure is ingested into the outboard tip vortex, as there is no significant modification to the progression of the OBVI

Aeronautical engineering: A continuing bibliography with indexes (supplement 291)

This bibliography lists 757 reports, articles, and other documents introduced into the NASA scientific and technical information system in May. 1993. Subject coverage includes: design, construction and testing of aircraft and aircraft engines; aircraft components, equipment, and systems; ground support systems; and theoretical and applied aspects of aerodynamics and general fluid dynamics

CFD analysis and design of a low-twist, hovering rotor equipped with trailing-edge flaps

This thesis reports the analysis and design of a hovering rotor equipped with both slotted and blended trailing-edge flaps. This was accomplished by combining a simple blade element method with 3D inviscid and RANS CFD that allowed for a robust sequence of design specification, analysis, and verification. Most modern helicopters have high levels of blade twist and various tip shape designs to help improve hover performance. However, such blade designs face problems due to compressibility effects on the advancing blade in forward flight. The twisted blade gives rise to negative incidence at the blade tip, which accelerates shock formation on the lower surface. The current work looks to evaluate the implementation of a low twist rotor for improved forward flight performance and recovering any potential losses in hover performance by deflecting fixed, trailing-edge flaps

Blade-Tip Vortex Noise Mitigation Traded-Off against Aerodynamic Design for Propellers of Future Electric Aircraft

We study noise generation at the blade tips of propellers designed for future electric aircraft propulsion and, furthermore, analyze the interrelationship between noise mitigation and aerodynamics improvement in terms of propeller geometric designs. Classical propellers with three or six blades and a conceptual propeller with three joined dual-blades are compared to understand the effects of blade tip vortices on the noise generation and aerodynamics. The dual blade of the conceptual propeller is constructed by joining the tips of two sub-blades. These propellers are designed to operate under the same freestream flow conditions and similar electric power consumption. The Improved Delayed Detached Eddy Simulation (IDDES) is adopted for the flow simulation to identify high-resolution time-dependent noise sources around the blade tips. The acoustic computations use a time-domain method based on the convective Ffowcs Williams–Hawkings (FW-H) equation. The thrust of the 3-blade conceptual propeller is\ua04%\ua0larger than the 3-blade classical propeller and\ua08%\ua0more than the 6-blade one, given that they have similar efficiencies. Blade tip vortices are found emitting broadband noise. Since the classical and conceptual 3-blade propellers have different geometries, especially at the blade tips, they introduce deviations in the vortex development. However, the differences are small regarding the broadband noise generation. As compared to the 6-blade classical propeller, both 3-blade propellers produce much larger noise. The reason is that the increased number of blades leads to the reduced strength of tip vortices. The findings indicate that the noise mitigation through the modification of the blade design and number can be traded-off by the changed aerodynamic performance