Showing papers on "Axial compressor published in 1968"

Improvements in or relating to blading arrangement for turbomachines

Abstract: 1,235,006. Turbines; axial flow compressors; jet propulsion plant. SOC NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION. Dec. 5, 1968 [Dec. 12, 1967], No. 57836/68. Headings F1C, F1J and F1T. A two-tier blade ring for a turbomachine such as a compressor ducted fan or the like comprises angularly spaced blades each formed by an inner section 2 having a tip portion 5 and a radially aligned outer section 3 having a base portion 6a mounted in a groove 6 in the tip 5. The blade sections 2, 3 are exposed to different working fluid flows, such as two air-flows or flows of air and combustion gases, and are separated by an annular member 1 having circumferentially spaced apertures engaging the tips 5 to strengthen the assembly. Insertion of each base 6a in its groove 6 is facilitated by a groove 7 in the member 1 in alignment with each aperture. As shown, the blades may be variable pitch having a spigot 4 secured in taper-roller bearings 8 in an annular member 9 fixed to the shaft r of a gas turbine engine compressor. Each spigot 4 carries Fig. 1, a gear 10 meshing with a rack 11 on an arm 12 secured to an axially slidable sleeve 13 mounted on shaft r. A hydraulic jack or other pneumatic, mechanical or electrical device can actuate the sleeve 13 to vary the blade pitch. In a modification, Fig. 6, the pitch varying mechanism comprises a toothed sector 21 secured to each spigot 4 and meshing with a gear 22 mounted on shaft r. Normally gear 22 rotates at the same speed as shaft r but can be rotated relative thereto to effect variation of the blade pitch. Fig. 5 (not shown) illustrates an arrangement in which two succesive two-tier compressor rotor stages have pitch varying mechanism as shown in Fig. 1, the respective control sleeves 13 being slidable in mutually opposite directions by a common hydraulic actuator (16) fixed to stator structure between the rotor stages. The two-tier stator blade rings (15) of the compressor may also be constructed similar to Fig. 1 with fixed or variable pitch blades.

Reynolds Number Effects in Axial Compressors

Abstract: The influence of variations in flow Reynolds number on the performance of axial compressors has been studied (changes in Reynolds number being, for the most part, achieved by changes in the inlet total pressure at or near the design speed). The measured results, so achieved, have been correlated to show how the main compressor performance parameters vary with Reynolds number. Reference has been made to cascade data to assist in choosing the form of the correlation, which is essentially empirical. A good correlation of the measured performance changes on component tests has been obtained. The method described, therefore, appears to be satisfactory for predicting trends for project assessments and avoids considering the detailed flow changes that occur within the machine as the Reynolds number is varied.

Gas turbine power plants having axialflow compressors incorporating contrarotating rotors

Abstract: 1,233,718. Gas turbine ducted fan engines; driving compressors. SOC. NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION. Dec. 11, 1968 [Dec. 14, 1967], No.59031/68. Headings F1C, F1G and F1J. A gas turbine engine comprises an axial-flow compressor and a turbine, the compressor being divided into n co-axial units where n is two or more, each compressor unit comprising two contra-rotating rotors, and the turbine comprises (n+1) independent rotors each driving a transmission shaft, the turbine rotating in opposite directions considered from one to the next, each turbine rotor driving two rotors of two successive compressors, except for the first and last turbine rotors which each drive only one compressor rotor. In the gas turbine ducted fan engine shown in Fig. 1 the axial flow compressor assembly 7 comprises an L. P. contra-rotating compressor 11 and an H.P. contra-rotating compressor 12 and the turbine assembly 9 comprises three independent turbine rotors 13, 14, 15 each driving a shaft 26, 27, 28. The engine comprises also a ducted fan 2 upstream of the L.P. compressor. The H. P. turbine 13 is connected by shaft 26 to drive the inner rotors 22, 24 of the H.P. compressor. The L. P. turbine 15 is connected by shaft 28 to drive the outer rotor 19, 21 of the L. P. compressor, also the ducted fan 2. The I, P, turbine 14 is connected by shaft 27 and the shaft 27a which are inter-connected by spline coupling 31, to drive the inner rotor 18, 20 of the L. P. compressor also the outer rotor 23, 25 of the H, P, compressor. The bearings for the shafts are indicated diagramatically. The ducted fan engine shown in Fig. 2 again comprises three turbine stages 13, 14, 15 and two contra-rotating compressor stages 11, 12 but in this case the H. P. compressor does not comprise a rotatable outer drum or casing. The H. P. turbine 13 drives through shaft 26 an internal ring gear 58 which drives a gear 56 which in turn drives a shaft 49 carrying gears 48 which engage internal ring gears 49 carried by the compressor discs 42 which in turn carry blades 40. The L. P turbine 15 drives through shaft 28 the outer rotor 19, 21 of the L. P. compressor also the ducted fan 2. The I. P. turbine 14 drives through shaft 27 the inner rotor 18, 20 of the L. P. compressor, also through gearing 59, 57 shaft 53 and gearing 52, 47 the discs 43 and blades 41 of the I. P. compressor. In the assembly three gears 56 engage with the internal ring gear 58 each gear 56 driving a shaft 49 and gears 48; similarly three gears 57 engage the internal ring gear 59 each gear 57 driving a shaft 53 and gears 52.

Gas turbine assembly having low-pressure groups and high-pressure groups adapted to be selectively connected either in series or in parallel

Abstract: A gas turbine installation which includes a combustion chamber, a compressor unit supplying the combustion chamber and having a low-pressure compressor and a high-pressure compressor, a drive turbine unit driven by the combustion gases and driving the compressor unit, and an output engine driven by the combustion gases, in particular an output turbine, whereby the low-pressure compressor and high-pressure compressor are adapted to be connected either in series or in parallel.

Axial flow compressor computer program for calculating off design performance /Program 4/

Abstract: Computer program for calculating off-design aerodynamic performance of axial flow compressor

Labyrinth seal for axial flow fluid machines

Abstract: A sealing device in a turbine for preventing excessive leakage of motive fluid between a stator and the tips of rotor blades, the leakage being from a high-pressure side to a low-pressure side, including an airfoil provided on the outer periphery of each rotor blade with the airfoil being skewed to pump the leakage fluid back to the high-pressure side.

Attenuation of circumferential inlet distortion in multistage axial compressors.

Abstract: Circumferential inlet distortion in a high hub-tip ratio multistage machine is treated by analyzing the compressor as a region in which a large number of small stages produce a pressure rise that is a function of the local mass flow rate. The resistance to circumferential flow due to the blading is included through an empirical factor. It is found that the over-all attenuation of both total pressure distortion and axial velocity distortion is mainly dependent on the slope of the compressor pressure rise vs flow rate characteristic. The attenuation increases when the slope of the characteristic is made more negative. In addition, considerable flow redistribution is found to occur upstream of the compressor. The theory has been compared with interstage data obtained on a three-stage, low-speed compressor with axial clearances that are 26% of the total length and a hub-tip ratio of 0.675. It is found that the approximation of zero axial clearance (infinite resistance to circumferential flow) gives excellent results. In consequence, it appears that for the normal range of axial clearances, the circumferential flow within the compressor can be neglected in a first-order analysis of the effects of inlet distortion.

The developing tangential velocity profile for axial flow in an annulus with a rotating inner cylinder

Abstract: A method is described of predicting the growth of a tangential velocity profile in fully developed laminar axial flow through a concentric annulus when the inner surface is rotated at speeds which are insufficient to generate Taylor vortices. The treatment, which is based on simplification and subsequent solution of the Navier-Stokes equations, as Fourier-Bessel series, appears preferable to momentum-integral techniques through greater simplicity of expression and in requiring fewer assumptions about the developing tangential profile. The validity of the predictions is best at high axial Reynolds number.

Aerodynamic evaluation of two-stage axial flow turbine designed for Brayton-cycle space power system

Abstract: Experimental design and performance test of two-stage axial-flow turbine designed for Brayton-cycle space power system

Variable Geometry Gas Turbine Radial Compressors

Abstract: An examination of the various mechanical and aerodynamic methods of flow and pressure regulation through high-speed, single-stage radial compressors is presented. Performance testing of a small gas turbine radial compressor with a variable vaned diffuser system is described. Methods of combined geometry variation to obtain minimum reduction of efficiency with flow variation are suggested based upon theoretical matching studies.Copyright © 1968 by ASME

Analysis of geometry and design point performance of axial flow turbines. 1 - Development of the analysis method and the loss coefficient correlation

Abstract: Stream-filament analysis procedure and correlation of total pressure loss coefficients to form basis of computer program to investigate design point performance of axial turbines

Experimental evaluation of outer casing blowing or bleeding of single stage axial flow compressor. Part 2 - Performance of plain casing insert configuration with undistorted inlet flow and boundary layer trip

Abstract: Outer casing blowing and bleeding and undistorted inlet flow for improved performance of high aspect ratio transonic rotor

Supercharger and fuel injector assembly for internal combustion engines

Abstract: The open air intake end of a hollow housing has adjustable air control shutters to regulate intake air to the transonic rotary vane component of an air compressor. The high velocity air output from the rotor is straightened to axial flow by passage through a stator vane component of the compressor at which combustion fuel is metered to the high velocity air. The fuel-air mixture then is expanded into a collector chamber from whence it is delivered to the intake valve system of an internal combustion engine.

Pressure Fluctuations Due to Interaction between Blade Rows in Axial Flow Compressors

Abstract: The paper presents the results of a theoretical and experimental investigation into the potential flow interaction effects between blade rows in an axial flow compressor stage. The investigation is concentrated on the inlet guide vane/rotor interaction and shows that the passing of the rotor blades behind the guide vanes produces large pressure fluctuations on the surfaces of the guide vanes.The available method of computation is not yet adequate for prediction of absolute values of pressure amplitude but does provide a sound basis for comparison between alternative designs.

Experimental evaluation of outer case blowing or bleeding of single stage axial flow compressor. Part 3 - Performance of blowing insert configuration no. 1

Abstract: Experimental evaluation of outer case blowing or bleeding of single stage axial flow compressor, and performance tests using distorted or undistorted inlet flow

Single stage experimental evaluation of slotted rotor and stator blading, part 9 Final report

Abstract: Slotted axial flow compressor rotor and slotted stator tests of slot effect on performance of highly loaded blades

Bladed rotor for a fluid flow machine

Abstract: 1,121,194. Compressor blades. ROLLSROYCE Ltd. May 1, 1967, No.20113/67. Heading F1T. Each rotor blade 20, particularly of the first stage of an axial compressor in a gas turbine engine, is provided intermediate its ends with projecting clapper portions 23, 24 which abut but are not secured to the respective clapper portions of the two adjacent blades in the rotor. The abutment faces 25, 26 are located in planes which are inclined in opposite senses to the axis of rotation of the rotor. The clapper portions are part-cylindrical in shape and may be integral with or secured to the blade.

Improvements in axial-flow turbines and compressors

Abstract: 1,099,677. Axial-flow compressors and turbines; jet propulsion engines. BRISTOL SIDDELEY ENGINES Ltd. Nov. 19, 1964 [Nov. 29, 1963], No.47274/63. Addition to 992,941. Headings F1C, F1J and FIT. In an axial-flow turbine or compressor for use with outlet ducting defined at least partly by a surface which is non-coaxial with respect to the rotor axis so as to be capable of causing a non- uniform distribution of the flow of working fluid through the rotor blades liable to cause damaging vibration of the rotor blades, at least a portion of at least one blade in a final ring of stator blades is pivotally mounted to permit individual adjustment of its exit angle, to enable the vibration to be reduced or eliminated. Fig. 4 shows part of a gas turbine engine, in which the exhaust duct 31 is branched to provide two lateral outlets 34 each having a rotatable gas deflecting nozzle 35. A bearing housing 33 is supported by a small number, say seven, of large vanes 32. To compensate for the variation of swirl in the exhaust gases leaving the last ring of turbine rotor blades 28 caused by variation in operation conditions of the engine, a row of variable-pitch stator blades 37 can be turned in unison in accordance with an operating condition of the engine, such as rotor speed, by means of a control ring 43 and arms 41. The stator blades 37 reduce the swirl angle to a value which can be dealt with satisfactorily by the relatively few vanes 32. The trunnion 40 of each stator blade 37 is adjustably attached to its arm 41 so that the blades 37 can be initially and individually adjusted in pitch whereby the pitch varies from blade to blade in such manner as to reduce or eliminate the tendency for the rotor blades 28 to be set into vibration by the non- uniformity of circumferential distribution of flow caused by the vanes 32, the branched shape of the exhaust duct 31 or the presence of the nozzles 35. Fig. 3 shows part of a by-pass type gas turbine engine in which individually-adjustable (but otherwise fixed) stator blades 19 are provided to compensate for the non-uniformity of circumferential distribution of flow in the by-pass annulus 16 tending to set the final rotor blades 10 of the lowpressure compressor into vibration. The non- uniformity may be caused by vanes 9a, which are few in number, say four, or by non-coaxial outlets. A square-section or splined extension 21 of the stem 20 of each blade 19 carries a vernier plate 22 seating on the casing 1, and the blade is locked by a nut 23 after adjustment. The blades 19, which preferably have no camber or twist, may be more or less equally distributed around the annulus 16 or may be provided in one or more groups each in a region where flow correction may be required. For example, they may be arranged in two arcuate groups at the top and bottom of the annulus 16 to compensate for the effect of an outlet opening at each side of the engine. Non-adjustable or adjustable stator blades 18 may be arranged with a circumferential variation of turning effect on the flow to offset the effect of flow disturbances caused by vanes 9b and, it may be, vanes 9a also. Adjustable flow-compensating stator blades provided at the rear of the compressor of a gas turbine engine may each have a non-adjustable leading edge part and an adjustable trailing edge part.