Showing papers on "Axial compressor published in 1983"

Pulsatile entrance flow in a curved pipe

Abstract: An experimental investitition was made of the entry flow in a curved pipe under conditions wherein a pulsatile component of flow was superimposed on a steady mean flow. Two experiments were conducted, one in a pipe of curvature ratio for αw = 12·5 and κ = 372. Laser-velocimetry measurements of the axial and secondary velocities were made throughout the cross-section at different instants of time within the cycle and at several axial locations downstream from the pipe entrance.The flow in the first experiment was found to be quasisteady, with axial and secondary velocities varying in time at the several axial stations proportionally to the instantaneous mean velocity, but essentially the same in character as low-Dean-number steady entry flow. The flow in the second experiment was more complex, with separation of the axial flow appearing at the inner bend during deceleration, starting downstream and propagating upstream toward the pipe entrance. Helical motions imbedded within the Dean circulation were also observed, and during certain portions of the cycle the secondary motions within the central core took on a jet-like structure.It suggested that the classifications employed by Smith to describe the different regimes of fully-developed pulsatile flow in curved pipes may also be useful to distinguish between different entry-flow regimes.

Development of Three-Dimensional Turbulent Boundary Layer in an Axially Rotating Pipe

Abstract: The time-mean velocities and turbulent fluctuations inside the turbulent boundary layers which developed in an axially rotating pipe were measured in the case where an undeveloped flow with a rectangular axial velocity distribution was introduced in the pipe. The pipe rotation gives two counter effects on the flow: one is a destabilizing effect due to a large shear caused by the rotating pipe wall and the other is a stabilizing effect due to the centrifugal force of the swirling velocity component of the flow. The destabilizing effect prevails in the inlet region, but the stabilizing effect becomes dominant in the downstream sections. The intensity of turbulence in the rotating pipe decreases ultimately below that in a stationary state of the pipe. Using the experimental results, the relationship between the mixing length and Richardson number proposed by Bradshaw was examined for the turbulent boundary layer that develops in the rotating pipe.

Back-skewed fan

Abstract: A compact, efficient axial flow fan with blades that are rearwardly (i.e., away from the direction of fan rotation) skewed and which are oriented at a pitch ratio which continuously decreases as a function of increasing blade radius along the radially outermost 30% of the blade.

A theory of rotating stall of multistage axial compressors

Abstract: A theoretical analysis was made of rotating stall in axial compressors of many stages, finding conditions for a permanent, straight-through traveling disturbance, with the steady compressor characteristic assumed known, and with simple lag processes ascribed to the flows in the inlet, blade passages, and exit regions. For weak disturbances, predicted stall propagation speeds agree well with experimental results. For a locally-parabolic compressor characteristic, an exact nonlinear solution is found and discussed. For deep stall, the stall-zone boundary is most abrupt at the trailing edge, as expected. When a complete characteristic having unstalling and reverse-flow features is adopted, limit cycles governed by a Lienard's equation are found. Analysis of these cycles yields predictions of recovery from rotating stall; a relaxation oscillation is found at some limiting flow coefficient, above which no solution exists. Recovery is apparently independent of lag processes in the blade passages, but instead depends on the lags originating in the inlet and exit flows, and also on the shape of the given characteristic diagram. Small external lags and tall diagrams favor early recovery. Implications for future research are discussed.

Pressure variation absorber

Abstract: Pressure variation absorbing apparatus is mounted adjacent to a moving fluid stream in the diffuser of a centrifugal compressor for absorbing both acoustic and aerodynamic pressure variations. The absorbing apparatus when mounted as a part of the diffuser wall of a centrifugal compressor not only reduces acoustic noise but also absorbs aerodynamic pressure variations increasing the efficiency of the compressor and simultaneously reducing the rate of flow at which surge occurs thereby enlarging the operational flow range of the compressor. A method of absorbing pressure variations in a moving fluid stream is also disclosed.

Secondary Flows and Losses in a Turbine Cascade

Abstract: : In designing axial flow turbines, one needs relatively simple prediction techniques for the angle variations and losses produced by secondary flows. To improve the physical basis of such techniques, an investigation was made of the flow in a cascade of large scale rotor blades of some 110 deg of turning. The aspect around 0.2 of the span. The flow was traversed in great detailed on ten planes using cobra-type probes with a computerised data recording and analysis system. Results are presented graphically using contour and vector plots on various planes through the flow field. The horse-show vortex and passage vortex development are clearly seen, with the upstream boundary layer being shed downstream as a loss core. Another region of high loss is related to a counter vortex in the corner between the sunction surface and the end wall. Results were also pitch averaged, and compared with predictions of angle and loss distribution. Some traversing was carried out using hot wire anemometry. Regions of high turbulence intensity and shear stress were identified, associated with the regions of high loss. Some indication is given of the mechanisms of loss production within the cascade.

Oil atomizing compressor working fluid cooling system for gas/vapor/helical screw rotary compressors

Abstract: Where a helical screw compressor gas or vapor working fluid is not too soluble in the compressor lubricant for compression process cooling, such lubricant after separation from the working fluid and cooled to as a low a temperature as possible and operating at as high a pressure as possible, is fed to an atomizing nozzle and injected into the inlet end of the compressor. This produces a cloud type blanket of cool atomized droplets uniformly dispersed within the suction vapor or gas allowing the highest possible rate of heat transfer to occur during the compression process and achieving the highest possible isothermal efficiency in a gas compression system or operation near vapor saturation values in a refrigeration vapor compression system while avoiding large hydraulic losses in the compressor. Working fluids such as helium, air, and ammonia may provide extremely high superheated gas or vapor at the compressor inlet.

Method for cylindrical grinding turbine engine rotor assemblies

Abstract: Blade tips of a turbine rotor assembly are ground to precision dimensions of ±0.025 mm (±0.001 inch) when the assembly is rotated at high speed during grinding. The speed should be sufficient to produce on the blades a radial acceleration of at least 700 times the force of gravity (6,900 m/s2). For compressor rotors the direction of rotation should be such that the blades move toward their concave side, and the speed should be sufficient to produce an axial flow. High resultant tip velocities are compensated for when dry grinding titanium alloys by moving the grinding media in the same direction as the tangential velocity at the point of contact. Burrs are eliminated by feeding at a very slow rate, of the order of 0.0004 mm/s when the final dimension is approached.

Three-dimensional structure and decay of vortices behind an axial flow rotating blade row.

Abstract: Detailed measurements were made of the three-dimensional turbulent flow field behind an axial-flow rotating blade row. The flow was surveyed at 15 radial locations and 70 circumferential sampling points in five measuring planes parallel to the trailing edge of the rotor. Statistically accurate mean velocities as well as turbulence stresses were obtained from numerous hot-wire signals, more than 12,000 for each sampling point. Vorticities were derived by the numerical differentiation of these data. The three-dimensional structure of various kinds of vortices generated through the rotor, such as a leakage vortex, trailing vortices, scraping vortices, a horseshoe vortex, etc. were elucidated quantitatively by use of the local streamwise, lateral and normal components of vorticity. The decay characteristics of these vortices were investigated in relation to the distribution of the turbulent stresses.

Gas turbine engines

Abstract: There is provided a high efficiency, medium output gas turbine engine which comprises at most three compressor stages, including one centrifugal compressor stage, all mounted on a common shaft for compressing inlet air, a combustion chamber for burning fuel in mixture with the compressed air to provide hot gases, output-power generating means operated responsive to the hot gases, an axial compressor-turbine driven by the hot gases exiting from the combustion chamber for driving an axial compressor integral with the axial compressor-turbine and constituting the final compressor stage, and a heat exchanger intercalated between the integral axial compressor/axial compressor-turbine unit and the output-power generating means. The heat exchanger is arranged to transfer heat from the hot gases leaving the axial compressor-turbine to the compressed air leaving the axial compressor. This abstract is not to be taken either as a complete exposition or as a limitation of the present invention, however, the full nature and extent of the invention being discernible only by referrence to and from the entire disclosure.

Control apparatus for centrifugal compressor

Abstract: A centrifugal compressor having a variable wall diffuser section that is controlled in response to measured system parameters in order to maintain the compressor at optimum efficiency over extended operating range.

The Effect of Inlet Distortion on the Performance Characteristics of a Centrifugal Compressor

Abstract: The present paper concerns itself with the effects of total pressure (and thus velocity) distortion on performance characteristics and surge margin of centrifugal compressors. Both radial and circumferential distortions were investigated. The performance tests as well as the velocity measurements within the impeller passages were carried out with a low speed compressor test rig with the inlet honeycomb as the distortion generators and compared with the case of “no distortion” as a datum. The results indicated that the inlet distortion exerted unfavorable influences on the efficiency and the surge margin of the given compressor, though the influence of the radial distortion was much stronger than that of the circumferential one. Various distortion indices were further examined in order to correlate the performance to the inlet distortion.Copyright © 1982 by ASME

Investigation of the Tip Clearance Flow Inside and at the Exit of a Compressor Rotor Passage—Part I: Mean Velocity Field

Abstract: A stationary two-sensor hot-wire probe has been used in combination with an ensemble averaging technique to measure the flow in the tip clearance region of a compressor rotor with emphasis on the leakage flow development inside the tip-clearance region and at the exit of the rotor. It is found that the presence and interaction of leakage flow, annulus boundary layer, and the scraping vortex is the dominant feature at mid-chord position. The rotation of the blade augments the leakage flow, resulting in the movement of the leakage jet toward the mid-passage. The blade-to-blade distribution of properties is shown to be highly nonuniform, except in the downstream and upstream regions.

Unsteady Three-Dimensional Turbine Aerodynamics

Abstract: High response aerodynamic measurements were made in a large scale, axial, flow turbine model to study the unsteadiness and three dimensionality of the flow. High response velocity vector and total pressure data were acquired. A comparison was made of the results of phase lock, averaging both raw and reduced data (voltages and velocities). The velocity vector measurements showed that there were strong radial flows present as well as significant periodic changes in the flow field due to relative rotor and vane positions. Random, periodic, and total unsteadiness levels were computed from the instantaneous and phase lock averaged velocity data. Time averaged data were compared with an inviscid two-dimensional calculation. A comparison was also made of time averaged total pressure measurements obtained from high response and low response (steady-state) probes.Copyright © 1982 by ASME

Axial-flow turbine blade, especially axial-flow turbine rotor blade for gas turbine engines

Abstract: The present axial-flow turbine blade has two cooling channels extending radially through the blade. These radial channels communicate with each other through impingement cooling ducts or holes (3) extending substantially axially through the blade. One channel (1) supplies cooling fluid through the holes (3) into the other channel (2) which is located nearer the leading edge (8) of the blade and extends substantially in parallel to the leading edge (8). The cooling fluid impacts on the wall of the other channel (2) and is laterally blown out at a side wall of the blade. For optimal cooling, especially of the leading edge of the blade, which is subject to severe thermal loads, this edge is penetrated by cooling elements (4) in the form of pins or wires in axial alignment with the impingement cooling holes or ducts (3). These cooling pins or wires (4) reach into the channel (2) nearer the leading edge with their tapered end (5) and are made of a highly thermally conductive material. The second cooling channel (2) nearer the leading edge (8) may have different cross-sectional shapes, for example the shape may be at least partially rounded along circumferential wall sections, or it may be circular, elliptical, elongated or the like for an improved guiding of the cooling fluid such as air.

Improvements in or relating to gas turbine engines

Abstract: A heat exchanged gas turbine engine power plant primarily intended for land vehicles has a two stage axial compressor (112) driven by a two stage low pressure turbine (114) through a shaft (116). A high pressure turbine (118) constitutes a power turbine which drives a load (not shown), through a shaft (122). The shafts are co-axial and can be coupled together by a free wheel mechanism (130). This arrangement of turbines in series in which the high pressure turbine drives the load enables the turbine entry temperature to be kept high over most of the power range without the need for variable power turbine inlet guide vanes. Also the engine provides a braking effect, tow starting is possible, and power turbine overspeed in event of load loss is checked. The heat exchanger ducting relatively simple and lightweight and the location of the combustion chamber (128) at the rear end of the engine makes access to the high temperature engine components relatively easy.

Medium bypass turbofan engine

Abstract: A medium bypass turbofan engine including a fan mounted on a low pressure shaft and positioned at the inlet of the engine; an intermediate pressure single-stage axial compressor carried by the low pressure shaft and positioned to receive a working airstream portion of the air discharged by the fan; a single-stage high pressure centrifugal compressor positioned to receive the output of the intermediate pressure compressor and carried by a high pressure shaft received telescopically over the low pressure shaft; an annular burner positioned to receive the output of the high pressure compressor; an annular slinger carried by the low pressure shaft and arranged to inject fuel into the burner; a high pressure single-stage axial turbine positioned to receive the output of the burner and carried by the high pressure shaft; and a low pressure multi-stage axial turbine positioned to receive the output of the high pressure turbine and carried by the low pressure shaft.

Stator assembly for an axial compressor

Abstract: A device for attaching the stator vanes of an axial compressor to an external housing is disclosed. The attachment utilizes an inner casing, having the vanes mounted on it, and an outer casing surrounding the inner casing in such a manner that differential thermal expansion between the various elements is minimized. This allows more accurate control of the clearances between the rotor and the stator.

Axial compressor middle stage secondary flow study

Abstract: This report describes an experimental investigation of the secondary flow within and aft of an axial compressor model with thick endwall boundary layers. The objective of the study was to obtain detailed aerodynamic and trace gas concentration traverse data aft of a well documented isolated rotor for the ultimate purpose of improving the design phases of compressor development based on an improved physical understanding of secondary flow. It was determined from the flow visualization, aerodynamic, and trace gas concentration results that the relative unloading of the midspan region of the airfoil inhibitied a fullspan separation at high loading preventing the massive radial displacement of the hub corner stall to the tip. Radial distribution of high and low total pressure fluid influenced the magnitude of the spanwise distribution of loss, such that, there was a general decreases in loss near the hub to the extent that for the least loaded case a negative loss (increase in total pressure) was observed. The ability to determine the spanwise distribution of blockage was demonstrated. Large blockage was present in the endwall regions due to the corner stall and tip leakage with little blockage in the core flow region. Hub blockage was found to increase rapidly with loading.

Variable capacity compressor and method of operating

Abstract: A plurality of unloading steps for a compressor (16) are achieved by providing the compressor with at least two cylinders (24, 25) having different displacements. By unloading different cylinders, the compressor output is better able to match system demand.

Measurements of reactive recirculating jet mixing in a combustor

Abstract: A ducted, subsonic, hydrogen/air turbulent jet mixing flowfield was experimentally investigated, both with and without combustion. The geometric configuration approximated a sudden expansion or "dump" combustor with a central air jet surrounded by a low-velocity hydrogen stream at an overall equivalence ratio of 0.12. The ratio of the duct to inner nozzle diameter was 2.5. Radial distributions of mean axial and radial velocity, axial and radial turbulent intensity, velocity cross correlation, gas composition, static temperature, and total pressure, as well as axial distributions of wall static pressure, were obtained for axial stations 0-6 duct diameters from the combustor entrance. The data indicate that mixing is slower in the chemically reactive flowfield than in the nonreactive flowfield, and that the presence of combustion has a significant effect on the size and location of the recirculation zone within the mixing duct.

Effects of Entrained Air on the Performance of a Horizontal Axial-Flow Pump

Abstract: The performance of an axial-flow pump was investigated under air-admitting conditions and the results were compared with those obtained for radial-flow pumps investigated in our previous studies. For the axial-flow pump we measured head as a function of impeller speed, liquid flow rate, and air void fraction. Flow visualization was performed and we correlated air bubble size as a function of Weber number based on mean impeller velocity. Overall flow patterns and the locations of air coalesence within the impeller were observed and diagrammed. These data contribute to earlier data obtained by ourselves and others to develop analytical models to predict pump performance degradation due to air admittance.

System for controlling the thrust nozzle adjustment of dual cycle gas turbine jet propulsion engines

Abstract: The thrust nozzle opening angle of two cycle gas turbine jet engines is crolled in a system which comprises a low pressure compressor driven by a low pressure turbine. A first compressor is arranged radially inward for a first flow cycle or circuit and a second compressor is arranged radially outward for a second flow cycle or circuit. A high pressure compressor is driven by a high pressure turbine. A combustion chamber is situated upstream of the high pressure turbine. An afterburner is supplied with the turbine gases from the first, hot flow cycle and with compressed air from the second, relatively cool flow cycle or circuit. The system may have multiple shafts. The control is effected with reference to the instantaneous pressure relationship (π NVZ ) between the static pressure (p sD ) downstream of the radially outer compressor (1a) of the second flow cycle (K2) and the static pressure (p sE ) of the air flowing into the propulsion system. This ratio constitutes a variable rated control value (π NVZ-Soll ) and is calculated as a function of the reduced rpm (N r ) of the low pressure compressor (1). This type of control guaranties during a steady state operation particularly a minimal fuel consumption in the subsonic region and a maximum thrust in the supersonic region. This type of control further assures during a nonsteady type of operating, particularly the stability between the two aerodynamic limits, compressor surges (G1 B or G1 H ) on the one hand and afterburner quenching (G2 B or G2 H ) on the other hand.

Stator Blade Row Geometry Modification Influence on Two-Stage, Axial-Flow Compressor Aerodynamic Performance.

Abstract: : The influence of stator row geometry modification on the aerodynamic performance of a two-stage, low-speed, axial-flow research compressor was assessed in the experiments described in this project. Stator geometry modifications tested included stator leading edge forward symmetrical sweep, large-radius blade/annulus wall corner fillets, and stator hub gap sealing (shrouding). Comparisons were made between detailed aerodynamic data associated with baseline and modified configurations. Substantial stator exit flow-field changes attributable to symmetrical sweeping of stator leading edges and to hub clearance sealing were observed with some evidence of corresponding near end wall loss reduction. The effects of large radii filleting were less clear. Interesting conclusions about the off-design flow rate performance of the compressor also resulted from consideration of experimental data. (Author)

Non-Uniform Flows in Axial Compressors Due to Tip Clearance Variation

Abstract: An approximate analysis is presented for the asymmetric velocity non-uniformities that can exist in an axial compressor due to variations in tip clearance around the circumference. Calculations have been carried out for several different compressor configurations, as well as for both rotating (fixed to rotor) and stationary clearance variations. The numerical results show that, for a given compressor, the ratio of (per cent) variation in axial velocity to (per cent) variation in pressure rise coefficient due to asymmetric clearance increases as flow decreases, although this trend in not necessarily true for the magnitude of actual velocity non-uniformity. In addition, the velocity non-uniformity due to a rotating clearance can be substantially larger than that due to a stationary clearance.