Rotating Fan Blades Research Articles

Analysis of Turbulent Boundary Layer on Cascade and Rotor Blades of Turbomachinery

The momentum integral technique for predicting the boundary-layer growth in three-dimensi onal flow has been extended to include the entrainment equation as the closure model. The numerical solution is compared with the cascade, inducer, compressor, and fan rotor blade data from various sources. The agreement is found to be excellent in all cases, with the exception of the separated flow. Both the momentum thickness and the limiting streamline angle predicted from this analysis compare well with the measured data for a rotor blade. The technique is extremely useful in engineering design, analysis, and performance prediction.

Characteristics of the Near Wake of a Compressor of a Fan Rotor Blade

Research reported in this paper covers an experimental investigation of the rotor wake. The experimental investigation included a study of the mean velocity, turbulence intensity, and Reynolds stress variations across the wake of a lightly loaded rotor blade at various axial and radial locations. Only the mean velocity data and their interpretation are presented in this paper. Measurements were carried out with a triaxial probe, rotating with the rotor and stationary behind the rotor. Also, measurements were made with a spherical head static pressure probe rotating with the rotor. Wakes were measured at various incidences to discern the effect of blade loading on the rotor wake. The wake is found to be three-dimensional in nature with appreciable radial velocity. The measurements close to the blade trailing edge indicate that the decay of the wake is rapid in this region. The wake data are correlated to derive expressions for the profile and decay of all the components.

Closure to “Discussion of ‘Optical Determination of Rotating Fan Blade Deflections’” (1977, ASME J. Eng. Power, 99, p. 210)

Closure to “Discussion of ‘Optical Determination of Rotating Fan Blade Deflections’” (1977, ASME J. Eng. Power, 99, p. 210)

Estimating the strength of turbine rotor blades taking into account operational damage

1. The law of equality of the action of opposite forces was used as the basis for determining the ratio of residual surface compressive stresses to subsurface tensile stresses in a hardened compressor rotor blade for the case of thin and thick cross sections. 2. Equations were derived which relate components of the stress state due to service loads to components of the initial stress field in the shaped portion of compressor and fan rotor blades. 3. An experimental and analytic method of determining the optimum level of compressive stresses in the hardened shaped portion of compressor rotor blades was developed and scientifically substantiated, the possible damage of the blade during its exploitation being taken into account in the development of the method. 4. Analysis of the problem showed that compressor and fan rotor blades must be thermally stabilized (before they are hardened by shot blasting) to produce complete relaxation of the nonuniform field of initial stresses induced in the shaped part of the blade during the mechanical treatment, since otherwise it is impossible to obtain a uniform initial stress field after the hardening treatment. 5. Compressor rotor blades hardened by shot blasting should be thermally stabilized at a temperature 10–15°C higher than the working temperature of the compressor blade at the given stage with a view to ensuring the optimum level of the uniform field of initial stresses in the shaped portion of the blade. 6. When actual rotor blades are fatigue tested to failure under laboratory conditions, their fatigue fracture will be analogous to that observed under actual working conditions only if the shaped part of the blade is subjected to the combined influence of alternating bending and an axial load simulating stresses due to centrifugal forces.

Month in the Patent Office

An aircraft powered by a gas‐turbine jet‐propulsion engine, the exhaust gases from which are discharged rearwardly through a jet pipe as a propulsive jet stream for normal forward flight, is provided with a fan mounted with its axis vertical, the fan drawing in air from atmosphere and discharging an air stream downwardly so as to produce an upward component of thrust on the aircraft, and means operable at will for diverting the exhaust gases from the jet pipe to drive the fan. The delta‐wing aircraft shown in FIG. 1 is powered by a gas‐turbine jet‐propulsion engine 3 which discharges exhaust gases as a propulsive jet stream through a jet pipe 5. An intermediate pipe 5a between the engine and the jet pipe is connected with two branch pipes 7 and is provided with a valve (not shown) which allows any desired proportion of the exhaust gases to be diverted into the branch pipes. The aircraft is provided with two fans, one in each wing, which rotate in opposite directions to balance out gyroscopic effects. As shown in FIG. 2, each fan rotor 12 comprises rotor blades 12b, a shroud ring 12c connecting the tips of the rotor blades, and a row of axial flow turbine rotor blades 12dmounted on the outer surface of the shroud ring. Exhaust gases diverted through the hand pipes 7 are supplied to these turbine rotor blades through a turbine inlet volute 13 which has a downwardly facing outlet provided with turbine nozzle vanes 14 co‐operating with the turbine rotor blades 12d. Rows of stator blades 19, 20 are provided above and below the fan rotor blades 12b, and these blades may be adjustable to enable the vertical lift to be varied; the pitch of the fan rotor blades may also be variable. Control of the aircraft may be effected by differential control of the blading of the two fans; alternatively or in addition the turbine nozzle vanes 14 may be adjustable. The inlets and outlets of the fans may communicate with the atmosphere through apertures in the wings which may be opened or closed by pivoted vanes 21, 22 which may be operated differentially to control the aircraft. Alternatively the inlets may be connected to one or more boundary layer suction openings in the surface of the aircraft. Additional fans 23, 24 may be mounted in the wing tips and nose of the aircraft to control the aircraft, these fans being driven by compressed air bled off from the compressor of the engine 3.

Month in the Patent Office

A helicopter rotor wherein each blade 3 carries a tip propulsion‐jet unit 4 has each unit constructed to reduce the velocity and noise level of the propulsive jets. Each unit 4 consists of an outer casing formed into a double wall at its forward end to provide an annular duct 12 which communicates through an aperture 20 with a conduit which extends through each blade from a gas producing plant 22, consisting of a gas turbine engine or engines or a separate compressor driven by such engine(s). Arms 10 secured radially on the forward end of the fairing support a central faired body 9 which is connected at its rear end to the fairing by stator blades 18 mounted ahead of a bladed rotor 5 secured in body 9 by a bearing 11. Blades 6 of rotor 5 carry a shroud ring formed on its periphery with turbine blades 8. Further stator blades 14, 15 connect the casing 4 to a tail fairing cone 16 and stator blades 13 bridge the duct 12. In operation gas flow from the blade conduit passes into duct 12 to be directed over blades 8 to drive rotor 5 which draws air through the unit to mix with the gases from gas producing plant 22 with a consequent reduction in jet velocity and noise. Any one of the rows of fan rotor and stator blades may be angularly adjustable to control the rotor speed. Alternatively, the turbine rotor blades 8 may be of the impulse type and a variable area admission nozzle used.