Showing papers on "Axial compressor published in 1989"

Laser anemometer measurements in a transonic axial-flow fan rotor

Abstract: Laser anemometer surveys were made of the 3-D flow field in NASA rotor 67, a low aspect ratio transonic axial-flow fan rotor. The test rotor has a tip relative Mach number of 1.38. The flowfield was surveyed at design speed at near peak efficiency and near stall operating conditions. Data is presented in the form of relative Mach number and relative flow angle distributions on surfaces of revolution at nine spanwise locations evenly spaced from hub to tip. At each spanwise location, data was acquired upstream, within, and downstream of the rotor. Aerodynamic performance measurements and detailed rotor blade and annulus geometry are also presented so that the experimental results can be used as a test case for 3-D turbomachinery flow analysis codes.

Redistribution of an inlet temperature distortion in an axial flow turbine stage

Abstract: The results of an experimental program aimed at determining the extent of the redistribution of an inlet temperature distortion in an axial flow turbine stage are presented. The program was conducted in a large-scale, low-speed, single-stage turbine where air, seeded with CO2, was introduced at one circumferential location upstream of the inlet guide vane. The migration of the seeded air through the turbine was determined by sensing CO2 concentration inside the stage. A temperature distortion was introduced by heating the seeded air. The CO2 concentration contours measured downstream of the vane showed h'ttle change with heating, indicating that the vane flowfield was relatively unaffected by the introduction of the temperature distortion. However, the CO2 contours observed on the rotor airfoil surfaces for the case with inlet heating indicated segregation of hot and cold gas with the higher temperature gas migrating to the pressure side and the lower temperature gas migrating to the suction side. Significant increases in rotor secondary flow were also observed. Two separate physical mechanisms are postulated to explain the observed experimental trends. Calculations performed by using a three-dimensional Euler solver show qualitative agreement with the experimental data.

Structure of Tip Clearance Flow in an Isolated Axial Compressor Rotor

Abstract: Ensemble-averaged and phase-locked flow patterns in various tip clearances of two axial compressor rotors were obtained by aperiodic multisampling technique with a hot wire in the clearance and with a high-response pressure sensor on the casing wall. A leakage flow region distinct from a throughflow region exists at every clearance. In the case of a small tip clearance, the leakage jet flow interacts violently with the throughflow near the leading edge, and a rolling-up leakage vortex decays downstream. As the clearance increases, a stronger leakage vortex comes into existence at a more downstream location, and a reverse flow due to the vortex grows noticeably. A scraping vortex is recognized at the pressure side near the trailing edge only for the small clearance. A horseshoe vortex appears in the upstream half of the through flow region for every tip clearance. The solidity does not affect the flow pattern substantially except for the interaction of the leakage vortex with the adjacent blade and wake.

Health Monitoring of Variable Geometry Gas Turbines for the Canadian Navy

Abstract: The Canadian Department of National Defence has identified a need for improved engine health monitoring procedures for the new Canadian Patrol Frigate (CPF). The CPR propulsion system includes two General Electric LM2500 gas turbines with multi-stage, variable-geometry axial compressors. A general method for predicting the thermodynamic performance of these compressors has been developed. The new technique is based on a meanline state-stacking analysis and relies only on the limited performance data typically made available by engine manufacturers. The method has been applied to the LM2500-30 marine gas turbine, and variations in engine performance that can result from a malfunction of the variable geometry system in service have been estimated.

Flow and Heat Transfer in Turbine Tip Gaps

Abstract: The effects of Reynolds number on flow through a tip gap are investigated by performing laminar flow calculations for an idealized two-dimensional tip gap geometry. The results of the calculations aid in understanding and reconciliation of low Much number turbine tip gap measurements, which range in tip gap Reynolds number from 100 to 10,000. For the higher Reynolds numbers, both the calculations and the measurements show a large separation off the sharp edge of the blade tip corner. For a high Reynolds number, fully turbulent flow calculations were also made. These also show a large separation and the results are compared with heat transfer measurements. At high Mach numbers, there are complex shock structures in the tip gap. These are modeled experimentally using a water table.

Application of Bifurcation Theory to Axial Flow Compressor Instability

Abstract: When a compression system goes unstable, the mode of response depends on the operating and system parameters, such as throttle setting and B parameter. Previous numerical work on the model developed by Moore and Greitzer (1986) has provided a limited picture of the parametric effects. Applying bifurcation theory to a single-harmonic version of the model has supplied much more complete information, defining the boundaries of each mode of response in the parameter space. Specifically, this is shown in a plot of B vs. throttle setting, which compares well with the corresponding map produced experimentally. The importance of the shape of the rotating stall characteristic is emphasized. The analysis shows a qualitative difference between classic surge and deep surge.

Stall Inception in Axial Compressors

Abstract: Detailed measurements have been made of the transient stalling process in an axial compressor stage. The stage is of high hub-casing ratio and stall is initiated in the rotor. If the rotor tip clearance is small stall inception occurs at the hub, but at clearances typical for a multistage compressor the inception is at the tip. The crucial quantity in both cases is the blockage caused by the endwall boundary layer. Prior to stall disturbances rotate around the inlet flow in sympathy with rotating variations in the endwall blockage; these can persist for some time prior to stall, rising and falling in amplitude before the final increase which occurs as the compressor stalls.Copyright © 1989 by ASME

Method and system for controlling variable compressor geometry

Abstract: A method and system for controlling the position of a variable pitch vane stage (34) in a gas turbine engine (10) uses a feedback loop comparing the current compressor pressure ratio to a reference compressor pressure ratio. The reference pressure ratio is determined from a plurality of current engine operating parameters which identify a unique point on the compressor operating line.

Variable stator vane arrangement for an axial flow compressor

Abstract: A variable stator vane arrangement for an axial flow compressor comprises a plurality of stages of variable pitch stator vanes each of which is operated by a respective control ring. The vanes in each stage are connected to the control rings by operating levers. An axially extending member is connected to each control ring by operating links. The axially extending member is arranged in a plane substantially tangential to each control ring and is movable in the plane substantially tangential to the control rings to change the pitch of the variable pitch stator vanes. Hydraulic rams are provided to move the axially extending member. The hydraulic rams may be operated to give either proportional or non-proportional movement of the variable pitch stator vanes. A controller may use signals from position detectors positioned on the control rings or position detectors positioned on the hydraulic rams together with compressor parameters to control the pitch of the vanes.

Weak rotating flow disturbances in a centrifugal compressor with a vaneless diffuser

Abstract: A theory is presented to predict the occurrence of weak rotating waves in a centrifugal compression system with a vaneless diffuser. As in a previous study of axial systems, an undisturbed performance characteristic is assumed known. Following an inviscid analysis of the diffuser flow, conditions for a neutral rotating disturbance are found. The solution is shown to have two branches; one with fast rotation, the other with very slow rotation. The slow branch includes a dense set of resonant solutions. The resonance is a feature of the diffuser flow, and therefore such disturbances must be expected at the various resonant flow coefficients regardless of the compressor characteristic. Slow solutions seem limited to flow coefficients less than about 0.3, where third and fourth harmonics appear. Fast waves seem limited to a first harmonic. These fast and slow waves are described, and effects of diffuser-wall convergence, backward blade angles, and partial recovery of exit velocity head are assessed.

Surge control in compressors

Abstract: Method and apparatus for operating compressors to avoid surges. The compressor is controlled by direct measurement of the rate of volumetric flow, Q, of gas through the compressor via the use of an acoustic flow meter, or other sonic velocity measuring device in combination with a standard flow meter and the measurement of the speed, N, of the compressor. The Q/N ratio is calculated electronically and output to a controller which, through a control circuit, operates a recycle valve in a by pass line leading from the discharge side to the suction side of the compressor. The Q/N ratio measured by the controller at compressor operating conditions is compared with a Q/N ratio at which surge will occur. When the measured Q/N ratio is less than the Q/N ratio at surge, the recycle valve will open and flow gas from the discharge side to the suction side of the compressor to avoid surge flow in the compressor.

Wake–Boundary Layer Interactions in an Axial Flow Turbine Rotor at Off-Design Conditions

Abstract: Interactions entre les sillages et les couches limites dans un rotor de turbine a ecoulement axial pour des conditions de fonctionnement non norminales

Part-Circumference Casing Treatment and the Effect on Compressor Stall

Abstract: Results are presented and discussed from an axial compressor rotor operated with an axial skewed slot casing treatment over part of the circumference. The compressor was one for which stall was initiated in the tip region and for this type there is some potential for stall margin improvement with lower loss using this. The main significance of the experiments is, however, the possibility of looking at aspects of stall inception. Normally stall inception is a brief transient with an unknown start time and is difficult to study but with the partial casing treatment it was possible to make the untreated section operate continuously in such a way that it underwent the processes normally leading to stall. For a tip stalling rotor the experiments identify the annulus boundary layer as the crucial region of the flow and spillage of the tip-clearance flow forward of the blades as a process leading to the rapid build up of blockage prior to instability and stall.Copyright © 1989 by ASME

Apparatus for venting the rotor structure of a compressor of a gas turbine power plant

Abstract: Apparatus for venting rotor parts of an axial flow compressor of a gas turbine power plant having an axial diffuser arranged behind the rotor and stator vanes at the last compressor stage. An annular chamber is formed downstream of the last rotor disk between adjacent rotor and stator parts and compressed air downstream of the diffuser is introduced into the chamber. The air in the chamber is divided into two individual streams, one of which is recycled to the compressor through an annular slot between the rotor and guide vanes at the last compressor stage and the other of which is discharged as a stream of leakage air into the secondary air system of the turbine.

J85 Surge Transient Simulation

Abstract: A one-dimensional model is presented for simulating surge disturbance propagations throughout an entire turbojet engine. The predicted results using this model are presented for the surge transient flow phenomena and the unsteady compressor operating-point excursion during surge in a J85 turbojet engine. The unknown stalled compressor stage characteristics are represented by parameters whose range was determined through a parametric study. The predicted stalled characteristics showed faster responses and larger pressure-ratio deterioration compared to low-speed compressors. Surge hammer overpressures were found to have a triangular shape in the time-space (engine axis) domain. The mass flow through the compressor was found to approach zero in 10-20 ms, thus affecting the peak hammer overpressures at the intake duct and the temperature rise in the compressor. Different stall methods were found to result in different compressor operating-point excursions.

Axial compressor blade assembly

Abstract: Blade platforms (26) for compressor blades with airfoils (22) in high solidity relationship have edges of a major portion (44) a minor portion (46), and a canted intermediate portion (48). Minimum clearance (54) exists between the major portions (44) whereby twisting during stackup or operation is avoided. A circumferential seal (60, 64) is located under the platform coincident with major portion (44).

Flutter of a Fan Blade in Supersonic Axial Flow

Abstract: An application of a simple aeroelastic model to an advanced supersonic axial flow fan is presented. Lane's cascade theory is used to determine the unsteady aerodynamic loads. Parametric studies are performed to determine the effects of mode coupling, Mach number, damping, pitching axis location, solidity, stagger angle, and mistuning. The results show that supersonic axial flow fan and compressor blades are susceptible to a strong torsional mode flutter having critical reduced velocities which can be less than one.

Numerical Results for Axial Flow Compressor Instability

Abstract: Using Cornell’s supercomputing facilities, we have carried out an extensive study of the Moore–Greitzer model, which gives accurate and reliable information about compressor instability. The bifurcation analysis in the companion paper shows the dependence of the mode of compressor response on the shape of the rotating stall characteristic. The numerical results verify and extend this with a more accurate representation of the characteristic. The effect of the parameters on the shape of the rotating stall characteristic is investigated, and it is found that the parameters with the strongest effects are the inlet length, and the shape of the compressor pressure rise versus mass flow diagram (i.e., tall diagrams versus shallow diagrams). We also discuss the effects of inlet guide vane loss on the characteristic. An evaluation is made of the h ′ = −g approximation, and a spectral analysis of the rotating stall cell given by the full model suggests why this breaks down.

Hoop shroud for the low pressure stage of a compressor

Abstract: An axial flow compressor for a bypass type jet engine has a low pressure stage which includes a one-piece, annular hoop shroud mounted to the airfoil of each of the rotor blades in the low pressure stage to divide the airflow entering the compressor into an outer, bypass airflow path and an inner airflow path directed to the higher pressure stages of the compressor. The hoop shroud is formed of circumferentially wound, parallel fibers such as silicone carbide fibers embedded in a metal matrix material such as titanium, wherein the volume ratio of the hoop shroud is approximately one-third fibers and two-thirds metal matrix material.

Subsonic axial flow fan noise and unsteady rotor force

Abstract: The noise radiated by a subsonic, axial flow fan at its rotational frequency and harmonics is related to the unsteady blade force created by the rotor blade/fluid interaction. This force is highly dependent on the time‐invariant flow distortions that enter the fan. The objectives of this study are to quantify the correlation between the inflow nonuniformity, unsteady blade loading, and the discrete‐frequency radiated noise. The inflow field of a typical cooling fan used in the electronic and computer industry was distorted by placing a small cylinder in the inlet plane. The fan was instrumented with a sensor to measure the axial rotor force. The on‐axis sound pressure spectra were compared to the coherent output power spectra conditioned on the unsteady rotor force and the radiated sound. Very good coherence at the discrete tones was observed. In the inlet plane, the time‐invariant, spatially nonuniform inflow was measured using miniature, five‐hole pressure probes. Fourier analysis of the inflow velocity data was used to compute the harmonic content of the unsteady rotor force. A simplified form of Curle’s equation was then used to compute the discrete‐frequency radiated noise at the blade passage frequency (BPF) and its lower harmonics. The predicted and measured noise levels are in close agreement at the BPF and the first harmonic, for which the fan is a compact source.

Process Plant Machinery

Abstract: Electric motors and controls gear speed transmission equipment gas turbines gas engines mechanical drive steam turbines turboexpanders centrifugal pumps positive displacement pumps vacuum pumps cooling water supply systems gas compression machinery axial flow compressors propeller, axial and centrifugal fans reciprocating compressors rotating positive displacement compressors mixers and agitators internal mixers - single and twin-screw extruders conveyor-based processing systems.

Hybrid thermal powered and engine powered automobile air conditioning system

Abstract: An air conditioning system with a common refrigerant supply, an engine-driven compressor circuit and a thermal compression circuit in parallel with the compressor circuit, a high efficiency compact thermal compressor which is either an ejector or a double vortex chamber flow inducer and a new control system which saves energy by unloading the mechanical compressor when thermal compression is fully operative. The hybrid system includes a check valve which prevents refrigerant flow from the compressor through the thermally powered compressor. Timers start the refrigerant pump of thermal compression circuit. Pressure switches sense operating pressure of the refrigerant. A refrigerant pump and a heat exchanger vapor generator generate high pressure refrigerant vapor which drives the thermal compressor. Refrigerant pump starts after initial cooling of passenger compartment is achieved by mechanical compressor circuit. Engine driven compressor is unloaded when outlet pressure of thermal compressor is high enough to drive system. THe precooling of the passenger compartment by mechanical compressor and subsequent unloading and deactivation of the engine-driven mechanical compressor permits lower fuel consumption. The system can operate with ozone-sparing refrigerants to replace CFC refrigerants, as well as with R12.

Unsteady Transition in an Axial Flow Turbine: Part 2 — Cascade Measurements and Modelling

Abstract: Part 1 of this paper reanalysed previously published measurements from the rotor of a low speed, single stage, axial flow turbine which highlighted the unsteady nature of the suction surface transition process. Part 2 investigates the significance of the wake jet and the unsteady frequency parameter. Supporting experiments carried out in a linear cascade with varying inlet turbulence are described together with a simple unsteady transition model explaining the features seen in the turbine.Copyright © 1989 by ASME