Showing papers on "Helicopter rotor published in 1969"

Unsteady aerodynamic and stall effects on helicopter rotor blade airfoil sections.

Abstract: Oscillatory tests in pitch and in vertical translation were performed on symmetrical and cambered airfoils, at full-scale Reynolds numbers, to provide dynamic stall data for rotor blade analyses. The Mach number range applicable to the retreating side of the rotor disk was covered at frequencies up to the first bending and the first torsional natural frequency. A system with a torsional degree of freedom was also tested. The following key results were found: 1) The negative aerodynamic damping due to stall is highly sensitive to Mach number. 2) Negative aerodynamic damping can be encountered in large-amplitu de plunging motions. 3) The maximum normal force encountered during oscillation is substantially higher than that for static stall. In addition, the flow separation process is discussed.

Rotor blade variable modulus trailing edge

Abstract: A helicopter rotor blade is provided with a trailing edge of hollow construction to form a cavity in which materials of various weights and densities is contained. The structural and aerodynamic properties of the rotor blade are enhanced by controlling the mixture of the modulus fibrous materials contained in the trailing edge cavity.

Articulated hub for helicopter rotor

Abstract: The present invention relates to a hollow, unit-construction mast-hub assembly having cylindrical arms at its top for supporting as many helicopter rotor blades through the medium of sleeves associated to said arms and pivotally mounted thereon, a flapping and lead-lag hinge device being interposed between each of said sleeves and a blade attachment clevis.

Studies of helicopter rotor noise

Abstract: : A comprehensive study of the problem of helicopter noise radiation is presented. After a review of the basic features of the noise, the limited experimental data are reviewed in some detail, and empirical laws are proposed. An exact theoretical expression for the noise is derived. This expression has been used as the basis for the development of a comprehensive computer program to calculate helicopter noise at any field point, including all effects of fluctuating airloads and all possible rigid and flexible blade motions. Under very reasonable approximations, an analytic expression was found for the sound field far from the helicopter, and computations based on this expression were made. The results show all the higher harmonics of the loading, at least up to the sixtieth, to be significant for noise generation. A study of the harmonic airloads is presented. Comprehensive acoustic results from the theory include the effect of various loading inputs, thrust, tip velocity, number of blades, blade motion, and forward speed. Fair agreement with experiment is found for overall levels, and good agreement is found with experimental trends. Design charts are presented which enable routine calculation to be made of noise radiated from any helicopter in hover or forward flight.

Variable geometry rotor system

Abstract: This rotor system is designed to control the nonuniform wake shed from a given rotor blade impinging upon the other blades of the rotor system. The rotor system utilizes blade sets which are of a different diameter than another blade set in the system. The azimuth spacing between the blade sets can be varied while the aircraft is in flight. The vertical spacing between the blade sets can also be changed. Mechanism is provided for collective pitch control of the blade sets. The plan form of blade sets, as well as the configuration of their tips, are varied.

Wind tunnel investigation of semirigid full-scale rotors operating at high advance ratios

Abstract: : A UH-1B 44-foot-diameter rotor having reduced-thickness tips was evaluated in a range of Mach numbers up to 0.94 and advance ratios of up to 0. 52. Additionally, UH-1D rotor blades reduced in diameter to 34 feet were tested at advance ratios of up to 1.1. Calculated performance is compared with the experimental results obtained to establish the validity of the theoretical technique at high advance ratios. In general, it was found that quasi-static, two-dimensional techniques were adequate up to an advance ratio of about 0.5. Above this advance ratio, theoretical techniques break down, especially with respect to calculating rotor propulsive force or drag. Theory-experiment comparison with the 44-foot-diameter rotor, operated at high Mach numbers, showed that Mach number effects are predictable to an advance ratio of at least 0.45. The 34-foot-diameter rotor became increasingly sensitive to control input with advance ratio. At an advance ratio of 1.1, this rotor system displayed a long transient response to a control input before obtaining its steady-state orientation, and at the largest values of collective pitch, the flapping would not completely stabilize.

A Correlation of Vortex Noise Data from Helicopter Main Rotors

Abstract: HE present and future use of helicopters as flexible air transportation in and out of highly restricted heavily populated areas is threatened by the community noise problems associated with these vehicles. In the next generation of vehicles, noise requirements may play a major role in the design process.1 It is important to have in the early stages of the design process an estimate for the effect of a given design change upon both the noise and DOC of the vehicle. This Note presents a correlation of existing data on main rotor vortex noise which is, at the present state-of-the-art, an irreducible effect of operating the main rotor. Although basic research into the effect of airfoil cross-section and tip planform shape on the noise from main rotors may eventually lead to noise reductions below this curve, an examination of the present data can also serve as a base to compare later improvements. For the typical helicopter, aerodynamic noise is produced by the main rotor, the tail rotor, and the engine. Current efforts in design should succeed in reducing the noise from the tail rotor and engine to a point where the dominant source of noise is the main rotor.

Control system for rotary wing vehicle

Abstract: A control system for a helicopter rotor has dual pitch changing mechanisms, one such mechanism for the inboard ends of the blades, and a second for servo flap control of an outboard station of each blade. Conventional cyclic and collective controls are provided in the vehicle, and the dual mechanisms operate in unison at moderate forward speeds, but function differentially at higher speeds to cyclically twist the rotor blades according to a predetermined airspeed schedule so that the advancing and retreating blades are cyclically twisted to reduce and to increase, respectively, the built-in negative twist of the torsionally resilient rotor blades.

Rotor blade pitch control by mechanical hydraulic system sensing blade deflection

Abstract: A system of mechanisms for automatically controlling, at relatively low speed, rotation of a helicopter rotor blade responsive to gust conditions or flapping loads or to reverse flow of air on the rear or trailing edge of a rotor blade. The system comprises novel blade sensor means, amplifying means coupled thereto, converting means, pitch azimuth sensing means, means for locking out gyroscopic control, and automatic low-speed pitch control means.

Rotor blade heavy core tuning weight

Abstract: A helicopter rotor blade includes a honeycomb core within the airfoil contour forming a composite-type rotor blade having the required mass for dynamically tuning the structure. The location and mass of the tuning weight is variable and the composite rotor blade is attachable to the hub in the conventional manner.

Isolation of Helicopter Rotor-Induced Vibrations Using Active Elements

Abstract: : Results of an analytical investigation of the feasibility of isolating helicopter fuselages from rotor-induced vertical vibratory forces while limiting the relative displacements during transient maneuvers and landing are presented Electrohydraulic elements are combined to provide better than 90 percent isolation at the critical rotor frequencies System parameters are selected for single-rotor helicopters ranging in weight from 2,000 to 80,000 pounds Results of the parametric study show the response of the electrohydraulic notch isolation systems to the various types of dynamic excitations in terms of rotor and fuselage transmitted accelerations, relative displacement between the rotor and fuselage, stability margin, power requirements, and estimated isolation system weight System performance and requirements are evaluated as a function of helicopter weight, blade passage frequency, number of notches of isolation, stability, changes in fuselage weight and rotor speed, and maximum allowable relative displacement during landing Recommendations are made regarding experimental verification of system performance, incorporation of approach into practical hardware, and isolation of combined vertical and in-plane rotor-induced vibrations

Prediction of rotor instability at high forward speeds. volume 3. stall flutter

Abstract: : The analytical study in this volume was carried out to determine the susceptibility of helicopter rotor blades to a stall flutter instability. This analysis was based on the use of unsteady aerodynamic data previously obtained for an NACA 0012 airfoil oscillating in pitch about its quarter-chord over a wide range of values of incidence angle, oscillatory frequency, amplitude of motion, and free-stream velocity. These data were originally available in the form of moment coefficient-incidence angle loops, and a twofold task was performed in carrying out this study. First, it was necessary to convert the moment coefficient data to an aerodynamic damping parameter form. This was accomplished by integrating the moment over one cycle of motion to yield the aerodynamic work per cycle, and this in turn was multiplied by appropriate conversion factors to produce the desired two-dimensional aerodynamic damping. Second, it was necessary to apply these two-dimensional results to a helicopter rotor to evaluate the weighted three-dimensional damping at each azimuth station, and to interpret the implications of any predicted region of instability. The stall flutter analysis was used in conjunction with the blade motion solution of Volume I to provide flight condition boundaries for stall flutter intensity.

Suppression of Transmitted Harmonic Vertical and Inplane Rotor Loads by Blade Pitch Control

Abstract: : A method was developed to study the possibility of using higher harmonic pitch-angle inputs to eliminate the transmission of oscillatory vertical and inplane forces from a helicopter rotor to its driving shaft. The aerodynamic loads are computed by using a realistic model which represents the rotor blades by bound vorticity distributions and the wake by a mesh of segmented vortex filaments. Computed results are presented for a two-bladed teetering rotor which was approximately the same as that of the UH-1A configuration except for the assumed differences in pitch control. The required pitch-angle inputs were determined for eliminating various combinations of harmonic root shears for three flight conditions.

Helicopter rotor construction

Abstract: A paired blade helicopter rotor having a strut connecting the blades and a tie rod connecting each blade with the rotor hub, provision being made to pretension the tie rods.

Whirl flutter of flapped blade rotor systems

Abstract: With a view to understand the whirl flutter phenomenon in flapped blade rotor systems and also to shed some light in bridging the existing gap between theoretical and experimental results, the equations of motion of an idealised mathematical model of a multi-bladed rotor system were formulated. Generalised aerodynamic forces were obtained from quasi-steady blade element theory. These13; equations were linearised and solved for a symmetric case of the system. The stability boundaries for different hinge offsets of the blades are presented as nacelle frequency parameter for neutral instability against blade flapping frequency parameter. It was found that for a particular value of the flapping frequency parameter, which is a function of blade hinge offset and restraint spring constant at the flapping hinge, the nacelle frequency parameter required for neutral instability is minimum. From these results it can be shown that an optimum flapping hinge position from the whirl flutter point of view can be determined. Low speed wind tunnel experiments were performed on a simplified model with different hinge offset conditions. Both 10 per cent and 13.6 per cent hinge offsets resulted in backward whirl flutter as predicted by theory.

Correlation of the vortex lattice method on rotor/wing configurations.

Abstract: The Hot Cycle Rotor/Wing is a new concept for high-speed vertical takeoff aircraft. It is a unique, dual-purpose lifting device that is basically a Hot Cycle helicopter rotor with an unusually large hub. It acts as a tip-jet-powered rotor for ver- tical and low-speed flight, and stops during flight to become a low-aspect-ratio, swept wing for high-speed cruise. By stopping the rotor in forward flight, the speed limitations of the helicopter rotor are removed, enabling more efficient cruise and operation at speeds up to 500 knots as a jet airplane. The single, dual-purpose lifting device combined with the simplicity and light weight of the Hot Cycle propulsion system holds promise of high payload capability superior to that of any other high-speed VTOL air- craft concept. Early low-speed wind-tunnel investigations2 of the concept considered the two planforms shown in Figs. 1 and 2 to the scale of the models. The tests of the circular and triangular hubs were conducted with the blades off and on in the stopped rotor mode. The reduced data include lift and pitching moment coefficients for angle of attack and rolling moment coefficient for antisymmetrical deflections of the left and right blades. Because the planforms are somewhat unconven- tional, methods of analysis for estimation of their aero- dynamic characteristics are rather limited. Perhaps the vortex lattice method of Hedman3 and the constant-pressure panel method of Woodward4 provide the only practical meth- ods of solution at this time. It is the purpose of this Note to present a correlation of the calculated results from Hed- man's method with the experimental data. Linearized estimates of the various measured coefficients are shown in Table 1. The reference area for the coefficients is the rotor disk area; the reference length for the moment coefficients is the rotor disk radius. The moments are taken about the disk center. Although many of the experimental

Helicopter rotor noise generation.

Abstract: Distributive loading rotational theory for helicopter rotor noise generation, considering steady and fluctuating force radiation and impulsive blade slap

Mechanism to transfer engine torque and control motion across helicopter rotor vibration isolator

Abstract: A helicopter rotor is attached to a structural member which is in turn mounted on an airframe by vibration isolators which enable the rotor and structural member to vibrate in a vertical plane and do not transmit these vibrations to the airframe. A linkage system having one bell crank pivoted on the structural member and another bell crank pivoted on the airframe is used to transfer control motion to the helicopter rotor with a minimum of interference from the vibration of the rotor and structural member. Universal joints are used to transfer engine torque from the engine mounted on the airframe to the vibrating rotor.

Improvements in or relating to the Construction of Rotor Blades for Rotary Wing Aircraft

Abstract: 1,150,123. Blades for rotary wing aircraft. WESTLAND AIRCRAFT Ltd. 11 Dec., 1967 [8 Feb., 1967], No. 6109/67. Heading B7W. A helicopter rotor blade 10 comprises a spar 11 which has root and tip portions 11a and 11b respectively of greater chord than the intermediate portion of the spar, the spar being of U-section. A shaped honeycomb core 12 is bonded inside the spar, and to spin plates 13, to complete the basic structure of the blade. The leading edges of plates 13 abut a protective nose capping 15 and the root portion is reinforced by external plates 20 formed with rotor head attachments 21. The nose portion of the blade carries a tube 16 for balance weights 17 and further balance weights may be located in tip portion 11b. Trim tabs 24 are provided.

The importance of vortex shedding effects on helicopter rotor noise with and without blade slap.

Abstract: Vortex shedding effects on helicopter rotor noise with and without blade slap, noting far and near field noise

Simplified procedures for estimating flapwise bending moments on helicopter rotor blades. Part 2 - Tables

Abstract: Helicopter rotor blades flapwise bending moments prediction by transfer function/superposition techniques

High Bypass Ratio Compound Fan-Shaft Engines for Convertible Rotary Wing Aircraft

Abstract: This paper presents design arrangements and performances of high bypass ratio compound fan-shaft engines. Helicopter rotor drive power can be obtained from the propulsion fan engine by unloading the fan. This is accomplished by using variable fan guide vanes that can be closed to reduce fan air flow or by reducing fan speed. These methods of unloading the fan are substantiated by the results of tests. The basic designs considered are for engines in the 5 to 30 lb/sec gas producer airflow size class and which have a two-spool coaxial front fan with a bypass ratio of approximately 8.0 and pressure ratio of approximately 1.4. The front fan configuration is used to superchange the gas producer. The flow of energy that results during power sharing between propulsive fan thrust and helicopter rotor drive power is discussed; in addition, fan shaft engine performance data are presented as applied to representative stowed-rotor convertiplanes of 15,000 and 60,000-lb design gross weights.Copyright © 1969 by ASME