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Investigation of rotor blade element airloads for a teetering rotor in the blade stall regime
A model of a teetering rotor was tested in a low speed wind tunnel. Blade element airloads measured on an articulated model rotor were compared with the teetering rotor and showed that the teetering rotor is subjected to less extensive flow separation. Retreating blade stall was studied. Results show that stall, under the influence of unsteady aerodynamic effects, consists of four separate stall events, each associated with a vortex shed from the leading edge and sweeping over the upper surface of the rotor blade. Current rotor performance prediction methodology was evaluated through computer simulation.
Forward sweep, low noise rotor blade
A forward-swept, low-noise rotor blade includes an inboard section, an aft-swept section and a forward-swept outboard section. The rotor blade reduces the noise of rotorcraft, including both standard helicopters and advanced systems such as tiltrotors. The primary noise reduction feature is the forward sweep of the planform over a large portion of the outer blade radius. The rotor blade also includes an aft-swept section. The purpose of the aft-swept region is to provide a partial balance to pitching moments produced by the outboard forward-swept portion of the blade. The rotor blade has a constant chord width; or has a chord width which decreases linearly along the entire blade span; or combines constant and decreasing chord widths, wherein the blade is of constant chord width from the blade root to a certain location on the rotor blade, then decreases linearly to the blade tip thereafter. The noise source showing maximum noise reduction is blade-vortex interaction (BVI) noise. Also reduced are thickness, noise, high speed impulsive noise, cabin vibration and loading noise.
An experimental investigation of the helicopter rotor blade element airloads on a model rotor in the blade stall regime
A wind tunnel test program was conducted on an eight foot diameter model rotor system to determine blade element airloads characteristics in the unstalled and stalled flight regimes. The fully articulated model rotor system utilized three blades with a Vertol 23010-1.58 airfoil section, the blades being 1/7.5 scale models of the Ch-47C rotor blades. Instrumentation was incorporated at the blade 75% radial station to measure pressure and skin friction distributions, surface streamline directions and local angle of attack. The test program was conducted in three phases; non-rotating, hover and forward flight at advance ratios of 0.15, 0.35 and 0.60. Test data were analyzed with respect to providing insight to the mechanisms affecting blade stall, particularly retreating blade stall during forward flight conditions. From such data, an assessment was made as to the applicability of current theoretical analyses used for the prediction of blade element airloads in the stall regime.
A Theoretical and Experimental Investigation of Flap-Lag Stability of Hingeless Helicopter Rotor Blades
The stability of hingeless rotor blade oscillations in hover is examined theoretically using a simplified centrally-hinged, spring-restrained, rigid blade to approximate the deflections of actual elastic blades. The aerodynamic and inertial coupling between the flap and lead-lag degrees of freedom is primarily responsible for instability, however elastic coupling and kinematic pitch-lag coupling both exert a powerful influence on hingeless rotor blade stability. Experimental results obtained from a two-bladed 1.81m diameter model rotor designed for minimum elastic coupling have confirmed the results of linear theory. For this model configuration rotor blade stall at high pitch angles was found to counteract the destabilizing flap-lag coupling and increase the damping of lead-lag oscillations. It was possible to account for this effect with the theory by using drag data for stalled airfoils.
A Dynamic Model Investigation of the Effect of a Sharp-Edge Vertical Gust on Blade Periodic Flapping Angles and Bending Moments of a Two-Blade Rotor
A two-blade rotor having a diameter of 4 feet and a solidity of 0.037 was subjected to sharp-edge vertical gusts while being operated at various forward speeds to study the effect of the gusts on the blade periodic bending moments and flapping angles. Variables studied included gust velocity, collective pitch angle, flapping hinge offset, and tip-speed ratio. Dimensionless coefficients are derived for the periodic components of the incremental changes in blade flapping angles and bending moments which arise when a rotor blade penetrates a sharp-edge gust. Mental changes in both the flapping angles and bending moments are essentially proportional to gust velocity, and the coefficients express the ratio of these increments to gust velccity. The results show that the flapping coefficient usually increases with an increase in collective pitch angle, is generally dependent on tip-speed ratio, and is essentially independent of the amount of flapping hinge offset. The bending-moment coefficient is also dependent on collective pitch angle and tip-speed ratio. Expected reductions in bending moments are realized by the use of flapping hinges, and further reductions in bending moments are achieved as the amount of flapping hinge offset is increased. Comparison of the experimental results of this investigation with limited available theoretical results shows substantial agreement but indicates that the assumption that the response of the rotor to a sharp-edge gust is independent of the collective pitch angle prior to gust entry is probably inadequate.
Performance Evaluation of Reduced-chord Rotor Blading as Applied to J73 Two-stage Turbine IV : Over-all Performance of First-stage Turbine with Reduced-chord Rotor Blades at Inlet Conditions of 35 Inches of Mercury Absolute and 700 R
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Experimental investigation of a high subsonic Mach number turbine having a 32-blade rotor with low suction-surface diffusion
High subsonic Mach number 32-blade rotor turbine with low suction-surface diffusion experiment to determine effect of low rotor blade solidity on losses
Experimental investigation of advanced concepts to increase turbine blade loading. 3 - Performance evaluation of tandem rotor blade
Performance evaluation of tandem rotor blade of single stage turbine
Optimizing tuning masses for helicopter rotor blade vibration reduction including computed airloads and comparison with test data
The development and validation of an optimization procedure to systematically place tuning masses along a rotor blade span to minimize vibratory loads are described. The masses and their corresponding locations are the design variables that are manipulated to reduce the harmonics of hub shear for a four-bladed rotor system without adding a large mass penalty. The procedure incorporates a comprehensive helicopter analysis to calculate the airloads. Predicting changes in airloads due to changes in design variables is an important feature of this research. The procedure was applied to a one-sixth, Mach-scaled rotor blade model to place three masses and then again to place six masses. In both cases the added mass was able to achieve significant reductions in the hub shear. In addition, the procedure was applied to place a single mass of fixed value on a blade model to reduce the hub shear for three flight conditions. The analytical results were compared to experimental data from a wind tunnel test performed in the Langley Transonic Dynamics Tunnel. The correlation of the mass location was good and the trend of the mass location with respect to flight speed was predicted fairly well. However, it was noted that the analysis was not entirely successful at predicting the absolute magnitudes of the fixed system loads.
Rotor blade system with reduced blade-vortex interaction noise
A rotor blade system with reduced blade-vortex interaction noise includes a plurality of tube members embedded in proximity to a tip of each rotor blade. The inlets of the tube members are arrayed at the leading edge of the blade slightly above the chord plane, while the outlets are arrayed at the blade tip face. Such a design rapidly diffuses the vorticity contained within the concentrated tip vortex because of enhanced flow mixing in the inner core, which prevents the development of a laminar core region.
Experimental investigation of advanced concepts to increase turbine blade loading. 2 - Performance evaluation of plain rotor blade
Performance of single stage turbine with rotor blade surface diffusion factor of 0.3
Forward sweep, low noise rotor blade
A forward-swept, low-noise rotor blade includes an inboard section, an aft-swept section, and a forward-swept outboard section. The rotor blade reduces the noise of rotorcraft, including both standard helicopters and advanced systems such as tiltrotors. The primary noise reduction feature is the forward sweep of the planform over a large portion of the outer blade radius. The rotor blade also includes an aft-swept section. The purpose of the aft-swept region is to provide a partial balance to pitching moments produced by the outboard forward-swept portion of the blade. The noise source showing maximum noise reduction is blade-vortex interaction (BVI) noise. Also reduced are thickness, noise, high speed impulsive noise, cabin vibration, and loading noise.
Simple theoretical models for composite rotor blades
The development of theoretical rotor blade structural models for designs based upon composite construction is discussed. Care was exercised to include a member of nonclassical effects that previous experience indicated would be potentially important to account for. A model, representative of the size of a main rotor blade, is analyzed in order to assess the importance of various influences. The findings of this model study suggest that for the slenderness and closed cell construction considered, the refinements are of little importance and a classical type theory is adequate. The potential of elastic tailoring is dramatically demonstrated, so the generality of arbitrary ply layup in the cell wall is needed to exploit this opportunity.
Effect of Rotor Blade Elasticity on UAM Quadrotor Acoustics
The 6-passenger quadrotor concept vehicle designed under the NASA Revolutionary Vertical Lift Technology (RVLT) Project is considered for acoustic analysis. The tip speed is 550 ft/sec, with three blades per rotor (550/3). Originally, the blades were rigid, uniform spanwise, and with flap and pitch degrees of freedom. The blade model has since been updated: a lag hinge was added, and nonuniformities and elastic properties were introduced. Four blade models are considered: 1) original model, rigid uniform flap-pitch; 2) rigid uniform flap-lag-pitch; 3) rigid nonuniform flap-lag-pitch; and 4) elastic nonuniform flap-lag-torsion. Predictions are made for three flight conditions (approach, flyover, and takeoff) using the four blade models. The RVLT Toolchain is exercised using CAMRAD II and pyaaron/AARON/ANOPP2. Quadrotor trim and performance, 0.75R vertical blade loading for all four rotors, and noise sources are analyzed. Also, the contributions of the front and rear rotor pairs to noise are studied. In approach and flyover, a 2 dBA loading noise difference (delta) is predicted between blade models 1 and 4 (delta for takeoff is smaller, 1 dBA). Most of this noise delta is due to the lag hinge and nonuniformities, which is consistent with the results of a 2022 study that had considered only the approach condition; the current results extend this conclusion to flyover and takeoff also. The insensitivity of quadrotor noise to blade elasticity is currently attributed to the high blade torsional stiffness (frequency 6.41 per rev) and the small blade radius (9 ft) of the 550/3 design. Suggestions for potential follow-on work are given.
Stiffness Characteristics of Composite Rotor Blades With Elastic Couplings
Recent studies on rotor aeroelastic response and stability have shown the beneficial effects of incorporating elastic couplings in composite rotor blades. However, none of these studies have clearly identified elastic coupling limits and the effects of elastic couplings on classical beam stiffnesses of representative rotor blades. Knowledge of these limits and effects would greatly enhance future aeroelastic studies involving composite rotor blades. The present study addresses these voids and provides a preliminary design database for investigators who may wish to study the effects of elastic couplings on representative blade designs. The results of the present study should provide a basis for estimating the potential benefits associated with incorporating elastic couplings without the need for first designing a blade cross section and then performing a cross-section analysis to obtain the required beam section properties as is customary in the usual one-dimensional beam-type approach.
Experimental investigation of advanced concepts to increase turbine blade loading. 4 - Performance evaluation of plain rotor blade with plow type vortex generators
Performance evaluation of plain rotor blade with plow type vortex generators
Vibration analysis of rotor blades with pendulum absorbers
A comprehensive vibration analysis of rotor blades with spherical pendulum absorbers is presented. Linearized equations of motion for small oscillations about the steady-state deflection of a spherical pendulum on elastic rotor blades undergoing coupled flapwise bending, chordwise bending, and torsional vibrations are obtained. A transmission matrix formulation is given to determine the natural vibrational characteristics of rotor blades with spherical or simple flapping pendulum absorbers. The natural frequencies and mode shapes of a hingeless rotor blade with a spherical pendulum are computed.