Computational Study of the Advancing-Side Lift-Phase Problem

During high-speed forward flight, helicopter rotor blades operate across a wide range of Reynolds and Mach numbers. Under such conditions, their aerodynamic performance is significantly influenced by dynamic stall—a complex... | Find, read and cite all the research you need on Tech Science Press

The innovative concept of asymmetric Helmholtz absorbers was recently introduced for applications in aerospace engineering. However, the influence of Mach number on the acoustic behavior of asymmetric Helmholtz absorbers remains underexplored. This study employs numerical simulations based on linearized Navier–Stokes equations to comprehensively investigate the acoustic properties of three-dimensional asymmetric absorbers under grazing flow. The analysis encompasses absorption and reflection coefficients as well as transmission loss, clarifying the complex relation between these parameters and the Mach number. The findings reveal that the maximum transmission loss of asymmetric absorbers is nonlinearly related to the Mach number, and a polynomial fitting function has been proposed to describe the law of maximum transmission loss as it relates to variations in Mach number and the spacing of the resonators. Specifically, in the Mach number range from 0.04 to 0.07, the transmission loss curve shifts from a single peak to two peaks. Sound incidence direction has little influence on the variation of reflection coefficient with increasing Mach number, but it does impact the absorption coefficient. Moreover, when the Mach number reaches a threshold value of 0.07, the influence of grazing flow on the interior neck becomes prominent, and the asymmetry disappears. The results of this study are expected to be useful for optimizing the design of asymmetric Helmholtz absorbers to enhance noise reduction in variable wind conditions.

A dynamic inflow model is used to calculate minimum induced power for a helicopter in high-speed forward flight with a finite number of blades. Comparisons between analytical and numerical results are shown, and they show good agreement. Different flow conditions (such as with and without reverse flow or inflow feedback) are used to show how each condition affects optimum induced power. Several results confirm the findings of earlier investigations, such as a singularity in rotor power in reverse flow and induced power reduction with an increase in blade number. One of the new findings is that the induced power is affected by rotor solidity (in the absence of inflow feedback) and the effect of inflow feedback does not always reduce the induced power. Results obtained using higher harmonic blade pitch control show that induced power can thereby be reduced for all conditions with a finite number of blades.

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, an in-house CFD solver based on Reynolds-Averaged NavierStokes (RANS). The acoustic analysis is performed using an acoustic code based on the Ffowcs William-Hawkings equation to solve for the noise propagating from the surfaces of the aircraft. 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 equivalent thrust between configurations, a 3 dB maximum OASPL reduction was obtained with the inclusion of the wing.

High-speed forward flight in helicopters causes high vibratory loads at the rotor hub, which are transmitted into the fuselage. This results in pilot fatigue and high maintenance requirements. Anti-vibration devices or vibration absorbers may be placed near the hub to reduce vibration transmission, while active vibration control systems may typically sense and reduce forces in specific areas of the fuselage. This paper focuses on developing strategies for optimal deployment (sizing, placement and control) of Circular Force Generators (CFGs) for vibration cancellation. Particle Swarm Optimization was used to carry out the optimization for two different load cases corresponding to different flight conditions. Converged solutions for CFG placement were obtained for: different individual load cases; multiple load cases; constrained and unconstrained actuator locations; and using different numbers of actuators. Two CFGs were found to be sufficient to cancel hub loads for a single load case. When two load cases are considered in multi-objective optimization, three actuators can substantially cancel the hub loads (to within 1%). And using four actuators can cancel the hub loads to within 0.1%. Finally, performance under actuator failure was considered as a way to distinguish among otherwise similarly-performing placement solutions.

This study characterizes the dynamics of a novel lag-pitch-coupled underactuated rotor design that can be incorporated into rotary-wing unmanned aerial vehicles (UAVs) to provide pitch and roll control with effectiveness comparable to that of a conventional swashplate albeit with significantly lower mechanical complexity and weight. The concept integrates a single lag hinge tilted at a 45-degree angle located at the center of the rotor hub with independent flap hinges for each of the two blades. This idea relies on the ability to cyclically vary the angular velocity of the rotor in a 1/rev fashion via motor torque modulation, which induces a cyclic lag resulting in a cyclic pitch variation due to the tilted lag hinge (lag-pitch coupling) and causes the tip path plane (TPP) to tilt in a desired direction for pitch and roll control. To understand this concept, simulations using the Rotorcraft Comprehensive Analysis System (RCAS) were performed to capture the 1/rev response in lag, pitch, flap and moments under steady flight conditions. The predictions were validated using hover and wind tunnel experiments. The results show that the maximum lag and pitch response are delayed by 180° relative to the maximum input torque. Additionally, Blade-1 and Blade-2 have different flap responses, which lag the peak pitch response by phase delays of approximately 15° and 40°, respectively. Wind tunnel data indicated an inherent coupling between pitch and roll moments as a function of the advance ratio. Finally, a single main rotor helicopter UAV was built and flight-tested to demonstrate the effectiveness of the concept in both hover and high-speed forward flight.

This study aims to employ the Lattice Boltzmann Method (LBM) to simulate the aerodynamic characteristics of intermeshingrotor and coaxial-rotor systems during high-speed and stable forward flight. The research entails a comparative analysis of rotor thrust coefficients and induced velocity distributions between these two rotor systems across various forward flight speeds. The primary objective is to investigate the feasibility of achieving high-speed performance with the synchropter and to gain valuable insights into the aerodynamic features of intermeshing rotors during high-speed and stable forward flight. The consistent findings of this investigation demonstrate that, under identical input parameters, the intermeshing-rotor system maintains superior thrust performance compared to the coaxial-rotor system. In forward flight, the interaction between the left and right rotors in the intermeshing-rotor system exhibits nearly identical characteristics, contrasting with the non-uniform interference observed between the upper and lower rotors in the coaxial-rotor system. Moreover, as flight speed increases, rotor interference in both systems tends to diminish. Notably, during high-speed and stable forward flight, the intermeshing-rotor system's performance surpasses that of the coaxial-rotor system, suggesting that the synchropter holds the potential to emerge as a novel configuration for high-speed rotorcraft.

This work describes the application of an adaptive control algorithm based on the generalized predictive control paradigmtoadetailednonlinearaeroelasticmodelofatiltrotoraircraftthatincludesrigid-bodydegreesoffreedom, restrictedtotheplaneofsymmetry.Theanalysisisbasedonanoriginalmultibodydynamicsformulation.Themodel is augmented by an autopilot and a stability augmentation system; in relevant cases, a model of the passive biomechanics of the pilot is considered as well. The interaction of rigid-body and deformable degrees of freedom alters the aeroelastic behavior of the system compared with that of the clamped wing rotor, discussed in a previous work. The flutter mechanism, originally consisting in whirl flutter, changes after considering rigid-body degrees of freedom;the short-period flight mechanics mode nowdominates high-speedstability. Thisissue ismainlyaddressed by the stability augmentation system, whereas the generalized predictive control is mainly used to reduce gustinducedwingloads,althoughitsinterventionforwhirl-fluttersuppressionisstillneededtoextendthe flightenvelope at very high speed. The capability of the adaptive regulator to reduce wing loads is assessed in hover and in highspeed forward flight. In the latter condition, the capability to suppress whirl flutter is assessed as well. HE use of active control in aerospace promises structural loads reduction and flutter suppression. However, setting aside all issues related to reliability, redundancy, and certification for a moment, modern aircraft (and rotorcraft) may not allow control designers to rely on the (sometimes abused) frequency separation paradigm that separately considers the dynamics of the aeroelastic and flight mechanics problems, thus significantly simplifying the analysis, since linearized models usually suffice for aeroelastic problems. In fact, airframe elastic modes with low frequencies interacting with low-frequency rotor modes may result in significant dynamic loading and reduced aeroelastic stability. This work focuses on the numerical analysis of the active aeroelastic control of a tiltrotor aircraft by means of an adaptive controller based on generalized predictive control (GPC), as discussed in [1], considering its interaction with flight mechanics dynamics, including the automatic flight control system (AFCS) and the passive dynamics of the pilot. Current generation tiltrotor aircraft mount two rotors on tilting nacelles that are located at the wing tips of an otherwise relatively conventional fixed wing aircraft design. Nacelle tilt changes the orientation of the rotor thrust about the pitch axis of the aircraft. As a consequence, tiltrotor aircraft can fly in two operating modes. In the helicopter mode (Fig. 1a) the nacelles are in the vertical position to provide the rotor thrust required to lift the aircraft. In helicopter mode, pitch control moments are generated by symmetrically tilting the rotor disks longitudinally, like conventional helicopters. Yaw control moment is produced in helicopter mode by differentially

This article investigates vibrational stabilization effects in rotorcraft flight dynamics. This study is motivated by recent results in flapping-wing flight, which showed that the time-varying aerodynamic and inertial loads due to the insect wing periodic motion induce a vibrational stabilization mechanism in hover. The dynamics of flapping-wing flyers and rotary-wing vehicles are both described by time-periodic systems so vibrational stabilization mechanisms can also have an effect on stability characteristics of rotary-wing vehicles. The article extends the use of the harmonic decomposition method to vibrational stability analysis of rotorcraft. Two cases are considered: vibrational stability due blade imbalance at hover, and vibrational stability due to number-of-blades-per-rotor-revolution (Nb\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$N_b$$\\end{document}/rev) in high-speed forward flight. Results show that while vibrations induced by rotor blade imbalance do not stabilize the hovering dynamics of a helicopter, these vibrations can still have a significant effect on the hovering dynamics. Rotor blade imbalance results in a symmetric effect on the roll and pitch axes, in that it tends to decrease the frequency of the subsidence modes of the hovering cubic, while the unstable oscillatory modes tend to increase in frequency and decrease in damping (destabilizing effect). On the other hand, the yaw/heave dynamics are relatively unaffected compared to the lateral and longitudinal axes. Moreover, Nb\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$N_b$$\\end{document}/rev rotor loads in forward flight are shown to reduce the damping of the coupled roll/pitch oscillation mode.

Limited available data suggest a substantial impact of Mach number on the heat transfer from an array of jets impinging on a surface at fixed Reynolds number. Many jet array heat transfer correlations currently in use are based on tests in which the jet Reynolds number was varied by varying the jet Mach number. Hence, this data may be inaccurate for high Mach numbers. Results from the present study are new and innovative because they separate the effects of jet Reynolds number and jet Mach number for the purposes of validating and improving correlations that are currently in use. The present study provides new data on the separate effects of Reynolds number and Mach number for an array of impinging jets in the form of discharge coefficients, local and spatially averaged Nusselt numbers, and local and spatially averaged recovery factors. The data are unique because data are given for impingement jet Mach numbers as high as 0.60 and impingement jet Reynolds numbers as high as 60,000, and because the effects of Reynolds number and Mach number are separated by providing data at constant Reynolds number because the Mach number is varied, and data at constant Mach number because the Reynolds number is varied. As such, the present data are given for experimental conditions not previously examined, which are outside the range of applicability of current correlations.

The operation of a propulsion system in terms of horizontal takeoff/landing and full-speed range serves as one of the main difficulties for hypersonic travelling. In the present work, a three-dimensional inward-turning inlet with tri-ducts for combined cycle engines is designed for the operation of three different modes controlled by a single rotational flap on the compression side, which efficiently simplifies the inlet structure and the flap control mechanism. At high flight speed between Mach 4 and 6, the pure scramjet mode is switched on, whereas both the ejector and the scramjet paths are open for a moderate Mach number between 2 and 4 with a larger throat area guaranteeing the inlet startability. In the low flight speed range with Mach number below 2, the additional turbojet path will be turned on to supply air for the turbine engine, whereas the other two paths remain open for spillage. Numerical simulations under different operation modes have proven the feasibility and good performance of the designed inlet, e.g., a nearly full mass flow ratio and a total pressure recovery around 0.5 can be achieved at the cruise speed. Meanwhile, the inlet works properly at low flight speeds which overcomes the typical starting problem of similar inlet designs. In the near future, wind tunnel experiments will be carried out to validate our inlet design and its performance.

A rotorcraft main-rotor blade experiences a broad range of the Mach number in high-speed forward flights or rapid maneuvering conditions. Near the boundary of the flight envelope, the rotor blade often encounters severe dynamic stall which limits the overall aerodynamic performance. Of interest here is effects of the Mach number on the dynamic stall. A rotor airfoil, VR-12, is computationally investigated for both static and dynamic stall conditions with varying the Mach number from 0.2 to 0.4. For the small enough Mach number 0.2, both static and dynamic stall are significantly influenced by flow separation due to adverse pressure gradient. For the highest Mach number 0.4, compressibility near the leading edge is no longer negligible, forming a shock there. The shock-induced separation occurs near the leading edge, which dominates both the static and dynamic stall at the high Mach number. At the intermediate Mach number 0.3, it is observed that both adverse pressure gradient and the shock affect the stall. Current computations are conducted with the unsteady Reynolds-averaged Navier Stokes approach with the Spalart-Allmaras model. Numerical results are compared to relevant wind-tunnel test data.

Oblique shock wave - boundary layer interactions (OSBLIs) at transitional Reynolds numbers are known to generate low-frequency unsteadiness that can impact aerodynamic efficiency, structural integrity, and system reliability in high-speed aerospace components. The present research examines how variations in Mach number, Reynolds number, and inviscid pressure jump influence the dynamics of these transitional interactions. Experimental campaigns were conducted in the TST-27 transonic–supersonic wind tunnel at TU Delft, utilizing high-speed and spark-light Schlieren imaging to capture flow behavior, followed by digital and spectral post-processing. The results reveal that transitional OSBLIs exhibit low-frequency shock oscillations strongly correlated with the periodic formation and disappearance of a Mach stem, a mechanism referred to as the “dual-domain” phenomenon. Systematic parameter variations showed that even small changes in Mach or Reynolds number substantially altered the presence of this dual domain as well as the amplitude and frequency of the oscillations. Furthermore, the Reynolds number regime previously identified as transitional for flat-plate boundary layers was experimentally validated in the present configuration. A secondary focus of the study was the application of passive flow control by introducing thin two-dimensional steps (≈60 μm) to trip the boundary layer. These perturbations were found to effectively suppress shock oscillations and modify the overall interaction dynamics, with frequency analyses confirming the disappearance of oscillation peaks observed in the baseline cases. A nondimensional analysis demonstrated a consistent Strouhal number convergence around St ≈ 0.33 for cases with strong unsteadiness. Increasing Reynolds number led to reduced laminar separation lengths and higher oscillation frequencies, consistent with expected transition dynamics. Overall, this research identifies the primary parameters governing unsteadiness in transitional OSBLIs, highlights effective suppression through minimal boundary-layer tripping, and provides nondimensional scaling to support predictive modeling. These insights contribute to improved understanding and design of aerospace components subject to shock-induced boundary-layer interactions.

With the recent surge of interest in manned hypersonic flight, most national and international hypersonic activities are focused on establishing a new hypersonic propulsion technology with the goal of providing more efficient access to space. This research is to demonstrate alternate propulsion technologies for access to space and high-speed flight within the atmosphere. As a preliminary step towards the construction and analysis of a wide class of airbreathing hypersonic vehicle, a FORTRAN code was developed. The development of this engineering tool is based on the elementary solution of the exact twodimensional, 2D, shock wave theory and on established empirical correlations that are reliable enough to predict very accurate information about the vehicle inviscid and viscous design characteristics. Further, the code is constructed in a very efficient manner, which allows for the design and analysis of a wide class of hypersonic systems, with an option to optimize the vehicle shape and aero-thermodymanic characteristics at the designer’s request. It is noteworthy to mention, that there are two important innovations to this design concept. First, a reliable engineering approach is developed to calculate the length of the scramjet combustor duct. This concept is based on the coupled solution of the onedimensional, 1D conservation of mass, momentum and energy equations with heat addition, combustion and friction. Secondly, a significant engineering design concept is introduced in the design of a minimum length nozzle, which is used as an integral part of the hypersonic vehicle configuration. This paper seeks to document this design methodology, and describes in detail the integration of the core elements as they relate to the overall hypersonic vehicle construction process. Finally, the code is used to conduct parametric studies on the behavior of the hypersonic configurations in relations to the vehicle flight conditions. In particular, the influence of the hypersonic vehicle lift, drag and aerodynamic efficiency are evaluated over a range of flight conditions. The flight conditions of interest to this study are mainly defined by the design parameters, such as, the Mach number, the flight altitude and the shock angles.

Direct numerical simulation of compressible turbulent channel flows using the discontinuous Galerkin method