Showing papers on "Blade pitch published in 1986"

Propeller module for an aero gas turbine engine

Abstract: A propeller module for a gas turbine engine comprising two contra-rotating propellers has a reduction gear positioned axially between the propellers. Pitch change mechanisms are provided to change the pitch of the blades of both propellers independent of the differential speed of the propellers.

Propeller/fan-pitch feathering apparatus

Abstract: A pitch feathering system for a gas turbine driven aircraft propeller having multiple variable pitch blades utilizes a counter-weight linked to the blades. The weight is constrained to move, when effecting a pitch change, only in a radial plane and about an axis which rotates about the propeller axis. The system includes a linkage allowing the weight to move through a larger angle than the associated pitch change of the blade.

Blade pitch varying mechanism

Abstract: An apparatus for varying the pitch of propulsor blades of a gas turbine engine by eccentrically driven gears. The gas turbine engine includes a rotating structure, an annular gas flowpath coaxial with the rotating structure, a plurality of rotor blades coupled to the rotating structure and extending into the gas flowpath such that a gas stream flowing through the flowpath causes the rotating structure to rotate with respect to the stationary member and a plurality of variable pitch propulsor blades coupled to and disposed radially outwardly of the rotating structure. The pitch of the blade is varied by a pinion gear coaxially coupled to one of the propulsor blades and an internal gear radially disposed about the pinion gear for driving the pinion gear. Annular displacement of the pinion gear causes angular displacement of the propulsor blade with respect to the rotating structure. The internal gear is revolved about the pinion gear whereby the pinion gear is angularly displaced with respect to the rotating structure.

Wind turbine pitch control hub

Abstract: In a variable pitch wind turbine system, pitch angle is determined by a pneumatically pressurized hydraulic actuator connected between crank arms on the blade pitch axles. The hydraulic line extends coaxially through the rotor drive shaft via a rotary union to a gas charged accumulator on the yaw carriage. The pitch axes are positively coupled for 1:1 counter-rotation by a gear train, preferably lying on the opposite side of the rotor axis from the hydraulic actuator.

An experimental investigation of the flap-lag-torsion aeroelastic stability of a small-scale hingeless helicopter rotor in hover

Abstract: A small scale, 1.92 m diam, torsionally soft, hingeless helicopter rotor was investigated in hover to determine isolated rotor stability characteristics. The two-bladed, untwisted rotor was tested on a rigid test stand at tip speeds up to 101 m/sec. The rotor mode of interest is the lightly damped lead-lag mode. The dimensionless lead-lag frequency of the mode is approximately 1.5 at the highest tip speed. The hub was designed to allow variation in precone, blade droop, pitch control stiffness, and blade pitch angle. Measurements of modal frequency and damping were obtained for several combinations of these hub parameters at several values of rotor speed. Steady blade bending moments were also measured. The lead-lag damping measurements were found to agree well with theoretical predictions for low values of blade pitch angle. The test data confirmed the predicted effects of precone, droop, and pitch control stiffness parameters on lead-lag damping. The correlation between theory and experiment was found to be poor for the mid-to-high range of pitch angles where the theory substantially overpredicted the experimental lead-lag damping. The poor correlation in the mid-to-high blade pitch angle range is attributed to low Reynolds number nonlinear aerodynamics effects not included in the theory. The experimental results also revealed an asymmetry in lead-lag damping between positive and negative thrust conditions.

Modular propeller blade pitch actuation system

Abstract: A modular, easily replaceable, pitch actuation system is removably secured within the hub of an aircraft. The modular system houses a driving system which provides a rotational output to set the pitch of propeller blades mounted to the hub, and a pitch changing system that controls the driving system. The driving system includes a high-speed, low-torque hydraulic motor that drives an harmonic drive to amplify the torque of the motor. The pitch changing system includes an actuator which positions a valve that controls the hydraulic motor, and a nulling gear which repositions the valve and acts as a pitch lock.

Nonlinear flexure and torsion of rotating beams, with application to helicopter rotor blades- i. formulation

Abstract: The differential equations of motion and their boundary conditions for a rotating beam such as a helicopter rotor blade are formulated via Hamilton's principle. The equations are va lid for both extensional and inextensional beams that have a pre-cone angle and a variable pitch angle. The equations are developed with the objective of retaining contributions due to higher order non-linearities which are generally disregarded in the literature due to their complexity. The influence of these higher order non­ linearities on the motion of a helicopter rotor blade is investigated by the authors in Part II of this paper that also appears in this issue of Vertica.

Experimental classical flutter results of a composite advanced turboprop model

Abstract: Experimental results are presented that show the effects of blade pitch angle and number of blades on classical flutter of a composite advanced turboprop (propfan) model An increase in the number of blades on the rotor or the blade pitch angle is destablizing which shows an aerodynamic coupling or cascade effect between blades The flutter came in suddenly and all blades vibrated at the same frequency but at different amplitudes and with a common predominant phase angle between consecutive blades This further indicates aerodynamic coupling between blades The flutter frequency was between the first two blade normal modes, signifying an aerodynamic coupling between the normal modes Flutter was observed at all blade pitch angles from small to large angles-of-attack of the blades A strong blade response occurred, for four blades at the two-per-revolution (2P) frequency, when the rotor speed was near the crossing of the flutter mode frequency and the 2P order line This is because the damping is low near the flutter condition and the interblade phase angle of the flutter mode and the 2P response are the same

High speed helicopter

Abstract: A high speed helicopter, ideally suited for use with rotors which are tip propelled, has an offset flapping hub mounted on a tiltable mast which is located a substantial distance off to the retreading blade side of the fuselage of the helicopter. This causes the fuselage to be located completely under the advancing blades. Individual pitch control motors individually vary the pitch of the rotor blades to control the feathering angle as a function of the angular position so that lift is obtained from the advancing and laterally moving blades only.

Concrete finishing machine having counterbalanced blade pitch adjustment apparatus.

Abstract: A concrete finishing machine having a rotatable trowel blade assembly (11) and a mechanism (27, 29, 33) for controllably adjusting the pitch of the trowel blades (15) relative to a wet concrete surface (19) on which the blades rest. Manual pivoting of a control lever (25) controllably adjusts the trowel blade pitch and, in doing so, automatically increases or decreases the amount of blade surface contacting the concrete surface (19), and correspondingly lowers or raises the machine relative to the surface (19). The contacting blade surface supports the machine's entire weight. A special counterbalancing apparatus that includes, for example, a compressed coil spring (51), biases the control lever (25) so as to compensate for the machine's weight and permits an operator to adjust the blade pitch with substantially less force than otherwise would be required.

Centrifugal fan with variable blade pitch

Abstract: A centrifugal fan includes a fan wheel having airfoil blades mounted for pivotal movement about their center of gravity. A yoke mechanism exterior of the fan wheel, when moved axially of the fan wheel drive shaft axis, causes actuator rods to move through the sidewall of the fan wheel. The axial movement of the actuator rods is translated first into rotary movement of a quill hub relative to the fan wheel and finally to radial movement of blade pivoting levers interior of the fan wheel. The movement of the blade pivoting levers causes the pitch of the individual fan blades to be changed in accordance with the degree of movement of the yoke mechanism. The yoke mechanism, actuator rods, quill hub and blade pivoting levers operate independently of the fan wheel so that blade pitch can be varied over an operating range irrespective of whether the fan wheel is stopped or is rotating. The yoke mechanism is moved and fan modulation is accomplished in response to the demand for conditioned air by the variable air volume system in which the fan is employed.

Aircraft thrust control

Abstract: An integrated control system for coaxial counterrotating aircraft propulsors driven by a common gas turbine engine. The system establishes an engine pressure ratio by control of fuel flow and uses the established pressure ratio to set propulsor speed. Propulsor speed is set by adjustment of blade pitch.

Application of higher harmonic blade feathering on the OH-6A helicopter for vibration reduction

Abstract: The design, implementation, and flight test results of higher harmonic blade feathering for vibration reduction on the OH-6A helicopter are described. The higher harmonic control (HHC) system superimposes fourth harmonic inputs upon the stationary swashplate. These inputs are transformed into 3P, 4P and 5P blade feathering angles. This results in modified blade loads and reduced fuselage vibrations. The primary elements of this adaptive vibration suppression system are: (1) acceleration transducers sensing the vibratory response of the fuselage; (2) a higher harmonic blade pitch actuator system; (3) a flightworthy microcomputer, incorporating the algorithm for reducing vibrations, and (4) a signal conditioning system, interfacing between the sensors, the microcomputer and the HHC actuators. The program consisted of three distinct phases. First, the HHC system was designed and implemented on the MDHC OH-6A helicopter. Then, the open loop, or manual controlled, flight tests were performed, and finally, the closed loop adaptive control system was tested. In 1983, one portion of the closed loop testing was performed, and in 1984, additional closed loop tests were conducted with improved software. With the HHC system engaged, the 4P pilot seat vibration levels were significantly lower than the baseline ON-6A levels. Moreover, the system did not adversely affect blade loads or helicopter performance. In conclusion, this successful proof of concept project demonstrated HHC to be a viable vibration suppression mechanism.

Helicopter blade and the like stand-off and folding device

Abstract: The instant invention provide a helicopter blade and the like stand-off and folding device that permits a helicopter blades to be folded up to positions close to the tail of the aircraft during inclement weather, storage, transport, or maintenance, etcetera. During a fold or unfold a one-to-one correspondence between the blades and the main rotor head is maintained and critical adjustments of the rotor blade assembly are not disturbed.

Development and testing of vortex generators for small horizontal axis wind turbines

Abstract: Vortex generators (VGs) for a small (32 ft diameter) horizontal axis wind turbine, the Carter Model 25, have been developed and tested. Arrays of VGs in a counterrotating arrangement were tested on the inbound half-span, outboard half-span, and on the entire blade. VG pairs had their centerlines spaced at a distance of 15% of blade chord, with a spanwise width of 10% of blade chord. Each VG had a length/height ratio of 4, with a height of between 0.5% and 1.0% of the blade chord. Tests were made with roughness strips to determine whether VGs alleviated the sensitivity of some turbines to an accumulation of bugs and dirt on the leading edge. Field test data showed that VGs increased power output up to 20% at wind speeds above 10 m/s with only a small (less than 4%) performance penalty at lower speeds. The VGs on the outboard span of the blade were more effective than those on inner sections. For the case of full span coverage, the energy yearly output increased almost 6% at a site with a mean wind speed of 16 mph. The VGs did reduce the performance loss caused by leading edge roughness. An increase in blade pitch angle has an effect on the power curve similar to the addition of VGs. VGs alleviate the sensitivity of wind turbine rotors to leading edge roughness caused by bugs and drift.

Tail rotor yaw position control for a helicopter

Abstract: A yaw rate stabilizer for the tail rotor of a helicopter includes a tail rotor shaft rotatably supporting a tail rotor on which rotor blades are radially mounted to develop a lift force directed substantially horizontal and transverse to the longitudinal axis of the aircraft. A slider is provided connected by pitch links to the rotor. The angle of attack of the blades varies with the displacement of the slider relative to a rotor. The rotor shaft is mounted pivotably by a universal joint on the tail of the helicopter fuselage. A frame, pivotably aligned with the universal joint, is supported by a tension or compression spring and connected to a collective pitch transfer lever. The slider, which is slidably mounted on the rotor shaft, moves toward the tail rotor and increases the collective pitch of the blades as a result of the rotor pivoting at the universal joint to precession caused by a transient, yaw moment applied to the rotor by a transverse gust of wind. The blades develop increased lift directed opposite to the gust which opposes the yaw moment and stabilizes the aircraft.

Aircraft propeller control

Abstract: In the invention, the speeds of both propellers in a counterrotating aircraft propeller pair are measured. Each speed is compared, using a feedback loop, with a demanded speed and, if actual speed does not equal demanded speed for either propeller, pitch of the proper propeller is changed in order to attain the demanded speed. A proportional/integral controller is used in the feedback loop. Further, phase of the propellers is measured and, if the phase does not equal a demanded phase, the speed of one propeller is changed, by changing pitch, until the proper phase is attained.

Controllable pitch aircraft propeller

Abstract: A variable pitch propeller, particularly suited for use with turbine engines, includes a range of beta control in low pitch and reverse pitch positions, in which the pilot can obtain a ground angle range with substantially lower angles in relation to the flight idle position. This is accomplished by the provision of lost motion connections in the form of sliding sleeves mounted on the beta feedback rods which come into operation at a particularly defined low pitch angle, under control of the pilot, and permits the blades to move to a still lower pitch angle for ground operations, by providing for relative movement between the blade yoke and the beta feedback rods.

Helicopter blade cyclic pitch control system

Abstract: A control and actuating system for adjusting the cyclic pitch of the blades of a helicopter automatically and independently in response to the blade gyrating angular position and to the helicopter forward velocity. This action is also independent of the control and adjustment of the collective pitch. The actuation of the blade pitch is performed by means of cams and followers. A hydraulic force-amplifying system is used between the follower/cam assemblies and the blade pitch actuator so as to isolate the follower/cam assemblies from the load carrying components of the system. The displacement signals generated by two cams are combined into a single linear displacement signal which causes a corresponding rotational movement of the blade. The axial position of one cam is determined by the pilot's action and the axial position of the other cam is defined by the helicopter forward speed. The two cams have surfaces that are not of revolution and which remain fixed for a given set combination of pilot's action and forward speed. The blade cyclic pitch variation is caused by the gyration of a housing to which the blades are attached and which contains and encloses the blade pitch control and actuating system.

Wind-driven ship

Abstract: Wind-driven ship, characterised in that it utilises wind energy from all directions for the ship propulsion by enabling sail, Flettner-rotor and wind-power-plant/propeller propulsion by one or more identical devices. The devices can be converted by virtue of the fact that the vanes of the vertical-axis-rotor wind-power plant are movable and either form a Flettner rotor when arranged on a centre diameter with their outsides or produce sails by stretching a sail material rolled up on the mast by connecting this sail material to two vanes which are run to the periphery. The rotatability of the central mast of the devices serves in following winds to orientate the sail, in side winds serves the rotary drive of the Flettner rotor by the ship's vertical-axis-rotor propeller driven by the relative movement of the water, and in head winds serves the rotary drive of the vertical-axis-rotor wind-power plant, which produces the ship propulsion via the ship's vertical-axis-rotor propeller. For stability reasons, to distribute the propulsion plant over several devices, for control reasons and to avoid flow losses by the ship's propeller, the devices are accommodated on the platform of a catamaran, and the ship's propeller can be retracted.

Device for controlling rotor blades

Abstract: The invention relates to a device for controlling rotor blades in the case of a helicopter, each rotor blade having at least one control unit which is connected to the rotor blade in such a manner that the longitudinal axis of said control unit runs parallel to the longitudinal axis of the rotor blade, it being possible for the control unit to carry out torsional movements about its longitudinal axis, without any mechanical drive.

Experimental study of the effects of installation on singleand counter-rotation propeller noise

Abstract: Measurements which are required to define the directivity and the level of propeller noise were studied. The noise radiation pattern for various single-rotation (SR) propeller and counter-rotation (CR) propeller installations were mapped. The measurements covered + or - 60 deg from the propeller disk plane and + or - 60 deg in the cross-stream direction. Configurations examined included SR and CR propellers at angle of attack and an SR pusher installation. The increases in noise that arise from an unsteady loading operation such as an SR pusher or a CR exceeded 15 dB in the forward axial direction. Most of the additional noise radiates in the axial directions for unsteady loading operations of both the SR pusher and the CR tractor.

Blade pitch measurement apparatus and method

Abstract: A blade pitch measurement apparatus comprises first means for providing an indication of the instantaneous angular position of the rotor relative to a reference location and second means for providing an indication of the passage of a tenon through a selected location about the rotor. A third means is operatively coupled to and responsive to the indications provided by the first and second means for providing an indication of the angular position of each tenon and for computing blade pitch. A method of measuring blade pitch is also disclosed.

Pusher propeller noise directivity and trends

Abstract: The effects of pylon wake interaction on far-field propeller noise are studied using a model scale SR-2 propeller in a low-speed anechoic wind tunnel. The variation in the pusher noise penalty with axial angle theta and circumferential angle phi is compared to that of the tractor noise penalty; and the former exhibits minima occurring in the propeller plane and maxima occurring toward the propeller axis. The magnitude of the pusher installation noise penalty decreased with in increase in shaft horsepower and tip Mach number. Directivity comparisons revealed that both a noise reduction and a directivity pattern change resulted when the pylon was moved farther from the propeller. Noise emerging from the wake interaction was distinguished from that of the propeller by means of a modal decomposition.

Bladed rotor assembly and a control system therefor

Abstract: A counterrotatable bladed rotor assembly drivable by a rotatable shaft comprises a leading propeller, a trailing propeller, and a gearbox located axially between the two propellers and enclosed by a non-rotatable casing having an end for attachment to a support structure. The gearbox has a first output shaft to the leading propeller and a second output shaft to the trailing propeller. A first fluid-pressure operable actuator is provided to vary the pitch of the blades of the leading propeller, and a second fluid-pressure operable actuator is provided to vary the pitch of the blades of the trailing propeller. In addition passageways are provided to connect the first and second actuators to a source of fluid under pressure. The casing rotatably supports the second output shaft at a position between the gearbox and the leading propeller and rotatably supports the first output shaft at a position between the gearbox and the trailing propeller. In this manner a compact and lightweight support arrangement is provided for the propellers.

Helicopter with high forward speed

Abstract: A helicopter having the usual tiltable, vertical thrust and lift producing rotor, a separate forward thrust producing propeller as well as provisions for yaw control and compensation is improved in that the propeller is constructed as unshrouded propulsion device. A pair of flow deflecting rudders is arranged downstream from the propeller and symmetrically to both sides of the craft; a frame on the tail portion of the craft includes vertical shafts for pivotally mounting the rudders. A T shaped support structure in the frame includes regular elevational, horizontally extending stabilizer fins and carrier arms. An elevator is arranged upstream from the propeller and from one of the elevational stabilizer fins. The elevator reaches laterally beyond and longitudinally alongside the one fin or beyond a circle provided by the propeller.

Main helicopter rotor

Abstract: A main helicopter rotor on which a rotary tubular shaft supports, at its top end, an annular hub integral with the tubular shaft itself. The assembly consisting of the tubular shaft and the hub is closed off at the top by a bell connected in removable manner to the hub and enclosing an oscillating plate assembly supported on a fixed shaft extending inside the aforementioned tubular shaft. A rotary ring on the oscillating plate assembly is connected to blade pitch change levers via respective rocker arms connected in articulated manner to the aforementioned rotary ring via respective pins extending radially in relation to the aforementioned bell and accessible via respective radial holes formed through the same.

Study of the propeller with small blades on the blade tips (1st Report)

Abstract: A new type propeller derived from “winglets” was investigated. This propeller is also fitted small blades at the blade tips likely to “winglets”, so we named it “bladelets”. These are for the purpose of reducing induced drag from tip vorteces. This paper, as the first report, mainly presents the results of the series tests which are concerned with the arrangement of these small blades. Propeller open test, flow visualization and flow velocity measurements were performed, and one of the best arrangement of “bladelets” were found out. The results looks 1 or 4% better compared the efficiency of “bladelet propeller” with that of the original one within the bound of real working condition.

A Study on the Three Bladed Propeller with Smaller Blade Area Ratio Designed by the New Method

Abstract: The authors designed prototypes of three bladed propellers with a smaller blade area ratio (expanded area ratio of 0.25) by a new design method that can take into consideration the variations of the effective attack angle in a wake. Model experiments were also carried out for the purpose of comparison between a newly designed propeller (SBA type) and a conventional propeller (MAU) type. The results at this time indicate that the new design method is useful for controlling and reducing propeller cavitation and hull surface pressure.