Showing papers on "Axial compressor published in 1984"

A Study on Configurations of Casing Treatment for Axial Flow Compressors

Abstract: Casing treatment is known to improve the stall margin of axial compressors. However, it is known as well that casing treatment lowers the efficiency of compressors. An experiment was planned in order to test a series of configurations of casing treatment which are supposed to have favourable effects and to obtain the optimum ones, if any, judged from above-mentioned two standpoints: larger stall margin improvement and smaller lowering of efficiency. The result shows that the relation between the stall margin imp improvement and the compressor efficiency for all of the tested configurations falls on some smooth curved-line, and thus, there exists no particularly superior treatment configuration to others. It rather shows that a certain amount of loss in efficiency is inevitable in order to obtain the required amount of stall margin improvement. Influence of rotor tip clearance on the effect of casing treatment was also examined. The result can be used as a guide for the selection of configurations in designing the casing treatment.

Theory of Blade Design for Large Deflections: Part I—Two-Dimensional Cascade

Abstract: As a step in the development of an analytical method for designing highly loaded, three-dimensional blade profiles for axial compressors and turbines, a simple two-dimensional method was first investigated. The fluid is assumed to be incompressible and inviscid, the blades of negligible thickness, and the mean tangential velocity is prescribed. The blades are represented by a distributed bound vorticity whose strength is determined by the prescribed tangential velocity. The velocity induced by the bound vortices is obtained by a conventional Biot-Savart method assuming a first approximation to the blade profile. Using the blade surface boundary condition, the profile is then obtained by iteration. It is shown that this procedure is successful even for large pitch-chord ratios and large deflections. In order to develop a method for use in three dimensions, the velocity is divided into a pitchwise mean value and a value varying periodically in the pitchwise direction. By using generalized functions to represent the bound vorticity and a Clebsch formulation for the periodic velocity, series expressions are obtained which can be adapted to three-dimensional problems. Several numerical results were obtained using both approaches.

Torque control apparatus for enclosed compressors

Abstract: A torque control apparatus for enclosed compressors is disclosed in which load torque applied to an enclosed compressor (13) is detected by a position detector (18) provided on the rotating portion of the compressor (13) and a temperature detector (17-1, 17-2) provided in a refrigeration cycle (13, 14, 15, 16), and a voltage or current supplied to the compressor (13) is controlled in accordance with the load torque to control the motor output torque of the compressor (13).

A Theory of Rotating Stall of Multistage Axial Compressors: Part II—Finite Disturbances

Abstract: An analysis is made of rotating stall in compressors of many stages, finding conditions under which a flow distortion can occur which is steady in a traveling reference frame, even though upstream total and downstream static pressure are constant. In the compressor, a pressure-rise hysteresis is assumed. Flow in entrance and exit ducts yield additional lags. These lags balance to give a formula for stall propagation speed. For small disturbances, it is required that the compressor characteristics be flat in the neighborhood of average flow coefficient. Results are compared with the experiments of Day and Cumpsty. If a compressor lag of about twice that due only to fluid inertia is used, predicted propagation speeds agree almost exactly with experimental values, taking into account changes of number of stages, stagger angle, row spacing, and number of stall zones. The agreement obtained gives encouragement for the extension of the theory to account for large amplitudes.

Theory of Blade Design for Large Deflections: Part II—Annular Cascades

Abstract: A method of designing highly loaded blades to give a specified distribution of swirl is presented. The method is based on a newly developed, three-dimensional analysis. In the present application, the flow is assumed to be incompressible and inviscid (the annulus has constant hub and tip radii), and the blades are of negligible thickness. A simple free vortex swirl schedule is assumed. The flow velocity is divided into circumferentially averaged and periodic terms. The Clebsch formulation for the periodic velocities is used, and the singularities are represented by periodic generalized functions so that solutions may be obtained in terms of eigenfunctions. The blade profile is determined iteratively from the blade boundary condition. Results from the computer program show how blade number, aspect, and hub-tip ratios affect the blade shape. The blade profiles for a given swirl schedule depend not only on the aspect ratio but also on the stacking position (i.e., the chordwise location at which this thin blade profile is radial), and so too do the mean axial and radial velocities. These effects occur whether the number of blades is large or small, and we conclude that even in incompressible flow the blade element or strip theory is not generally satisfactory for the design of high-deflection blades. The analysis derives the geometrical conditions for the blade profiles on the walls of the annulus which are needed to satisfy the wall boundary conditions in the idealized flow, but which in any practical example will be modified by the presence of wall boundary layers and blade thickness. In the limit when the number of blades approaches infinity, a bladed actuator duct solution is obtained. The conditions for the blade profile at the walls are absent, but the stacking position and aspect ratio still affect the axial and radial velocity distributions for the same swirl schedule.

Experimental Study of a High-Throughflow Transonic Axial Compressor Stage

Abstract: Design information and experimental results are presented for a transonic axial compressor stage passing 40 lbs/s-ft/sup 2/ frontal area (195 Kg/s-m/sup 2/) with a pressure ratio of 1.95 at 1500 ft/s (457 m/s) tip speed. The design incorporates several unusual features that helped it achieve a peak isentropic efficiency over 88 percent at design speed. The compressor was evaluated at three rotor tip clearances and an optimum was found. Vortex generators placed upstream on the casing proved relatively ineffective in influencing stall margin. Vortex generators installed on the rotor did improve stall margin and also increased efficiency at speeds of 90 percent and below.

Performance of a high-work low aspect ratio turbine tested with a realistic inlet radial temperature profile

Abstract: Experimental results are presented for a 0.767 scale model of the first stage of a two-stage turbine designed for a high by-pass ratio engine. The turbine was tested with both uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The inlet temperature profile was essentially mixed-out in the rotor. There was also substantial underturning of the exit flow at the mean diameter. Both of these effects were attributed to strong secondary flows in the rotor blading. There were no significant differences in the stage performance with either inlet condition when differences in tip clearance were considered. Performance was very close to design intent in both cases.

Active clearance control

Abstract: The bore of the compressor for a gas turbine engine is heated by selectively bleeding compressor air from downstream stages so that heating only occurs at discreet times during the engine operating envelope. The bled air is admitted into the bore at the mid-stage station of the compressor wherein the compressor disks are scrubbed so as to expand and close the gap between the outer air seal and tips of the compressor blades during cruise of the aircraft and prevented from heating the disks during the high powered operations of the engine.

Principles of turbomachinery

Abstract: List of symbols used, their meaning and dimensions. Fundamental principles. Principles and practice of scaling laws. Principles of axial flow machines. Principles of radial and mixed flow machines. Centrifugal machines. Axial machines for incompressible flow. Axial turbines and compressors for compressible flow. Radial flow turbines. Cavitation and other matters.

Bleed valve for axial flow compressor

Abstract: A bleed valve construction for a multistage axial flow compressor in which a series of bleed doors in frames located between adjacent struts of a support case between successive stages of the compressor are actuated by a ring that has only circumferential movement imparted by a rack on the ring driven by a motor supported pinion.

Experimental cascade analysis of a transonic compressor rotor blade section

Abstract: A transonic compressor rotor blade cascade was tested in order to elucidate the flow behavior in the transonic regime and to determine the performance characteristic in the whole operating range of a rotor blade section. The experiments have been conducted in a transonic cascade wind tunnel, which enables tests even at sonic inlet velocities. The influence of the upstream Mach number between 0.8 and 1.1 and the inlet flow angle between choking and stalling of the blade row was investigated. The effect of the axial velocity density ratio (AVDR) could be studied by applying an endwall suction device. Furthermore, the level of the shock losses was determined from a wake analysis. A final comparison of cascade losses and those of the corresponding rotor blade element shows good agreement which underlines the applicability of the cascade model in the design of axial flow turbomachines.

Redesign and performance analysis of a transonic axial compressor stator and equivalent plane cascades with subsonic controlled diffusion blades

Abstract: In order to verify a new controlled diffusion blade design concept, the stator of an existing transonic axial compressor stage was redesigned. Stator and equivalent cascade tests revealed the potential of such blades for a considerably higher aerodynamic loading than it has been applied up to now. The design procedure is described, and the results of plane cascade and stage testing are submitted, including performance analysis of both cascade and stator blade sections, at design and off-design operating conditions. The blade profile shapes and cascade geometries were calculated by means of an inverse, two-dimensional method taking also into account the axial velocity density ratio (AVDR). This design concept is essentially based on prescribed blade pressure distributions, which are optimized with respect to the blade boundary layer development. The flow phenomena are illustrated by means of loss and flow turning investigations, blade pressure distributions, and laser velocimetry data. The test results reveal that the two-dimensional approach applied is quite promising for the three-dimensional blade design. Finally, overall and blade element performance comparisons between the original NACA 65 profiled stator and the redesigned one demonstrate the favorable flow behavior of the new stator, as well as the great potential of the controlled diffusion blade concept.

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.

Extended parametric representation of compressor fans and turbines. Volume 1: CMGEN user's manual

Abstract: A modeling technique for fans, boosters, and compressors has been developed which will enable the user to obtain consistent and rapid off-design performance from design point input. The fans and compressors are assumed to be multi-stage machines incorporating front variable stators. The boosters are assumed to be fixed geometry machines. The modeling technique has been incorporated into time sharing program to facilitate its use. Because this report contains a description of the input output data, values of typical inputs, and examples cases, it is suitable as a user's manual. This report is the first of a three volume set describing the parametric representation of compressors, fans, and turbines. The titles of the three volumes are as follows: (1) Volume 1 CMGEN USER's Manual (Parametric Compressor Generator); (2) Volume 2 PART USER's Manual (parametric Turbine); (3) Volume 3 MODFAN USER's Manual (Parametric Modulating Flow Fan).

A Three-Dimensional Model for the Prediction of Shock Losses in Compressor Blade Rows

Abstract: A design trend evident in newly evolving aircraft turbine engines is a reduction in the aspect ratio of blading employed in fans, compressors, and turbines. As aspect ratio is reduced, various three-dimensional flow effects become significant which at higher aspect ratios could safely be neglected. This paper presents a new model for predicting the shock loss through a transonic or supersonic compressor blade row operating at peak efficiency. It differs from the classical Miller-Lewis-Hartmann normal shock model by taking into account the spanwise obliquity of the shock surface due to leading-edge sweep, blade twist, and solidity variation. The model is evaluated in combination with three test cases. Each was a low-aspect-ratio transonic stage which had exceeded its efficiency goals. Use of the revised shock loss model contributed 2.11 points to the efficiency of the first test case, 1.08 points to the efficiency of the second, and 1.38 points to the efficiency of the third.

Stalled Flow Performance for Axial Compressors: I — Axisymmetric Characteristic

Abstract: A study of stalled flow performance of multistage compressors is presented. A new compressor characteristic is developed, describing the axisymmetric pumping performance over the entire compressor flow range, including reversed flow. This axisymmetric characteristic is required in any current rotating stall model. It is possible for the axisymmetric performance to rise above the measured stall point pressure rise, thus indicating greater unstalled pressure rise potential. In this context, the axisymmetric characteristic in forward flow is viewed as paralleling diffuser performance. A simple two-dimensional reversed flow model is presented, and is shown to be in reasonable agreement with available high backflow compressor data. The model predicts that the blade stagger angle greatly influences the reversed flow characteristic. Calculations are carried out using the rotating stall model of Moore and the axisymmetric characteristic developed herein, and a technique is suggested for estimating the axisymmetric curve over the entire flow range.Copyright © 1984 by ASME

Preventing surge in an axial flow compressor

Abstract: In axial flow compressors of gas turbine engines it is well known to use either an air bleed valve 34 or a compressor casing treatment to improve the performance of the compressor. Air bleed valves act to keep the working line of the compressor close to the surge line. The casing treatment, which usually consists of one or more grooves 28 in the internal cylindrical surfaces adjacent the tips of a stage of upstream blades 22, acts to raise the compression ratio at which the rotor stalls. This invention combines the use of the bleed valve 34 with the grooves 28 to allow the bleed valve setting point to be reduced and ensure the bleed valve is closed at low power settings.

Compressor casing recess

Abstract: A means for improving the aerodynamic efficiency of the compressor of an axial flow turbomachine is disclosed. The compressor includes an airfoil relatively rotatable with respect to a radially disposed surface which bounds a flowpath for aft moving fluid. The surface has a circumferentially extending recess radially disposed relative to the airfoil. The recess has a generally aft facing wall and a generally forward facing wall. The aft facing wall is oriented so as to provide a barrier to the forward flow of fluid in the clearance between airfoil and surface. The forward facing wall is oriented so as to provide an aerodynamically smooth transition from the recess into the flowpath.

Mass Flow Limitation of Supersonic Blade Rows due to Leading Edge Blockage

Abstract: The supersonic flow passing through the blade row of an axial flow compressor depends on the magnitude of the axial inlet Mach number. If the upstream stream tube is convergent, the axial Mach number remains subsonic and the unique incidence relation holds between inlet Mach number and inlet flow angle. This paper presents the influence of leading edge thickness and suction surface angle on the unique incidence and choking of supersonic blade rows. The theoretical results were experimentally verified in supersonic cascade wind tunnel tests.Copyright © 1984 by ASME

Compressor casing recess

Abstract: A means for improving the aerodynamic efficiency of the compressor of an axial flow turbomachine is disclosed. The compressor includes a first airfoil relatively rotatable with respect to a radially disposed surface and a second airfoil, aft of the first airfoil, and fixed with respect to the surface. The surface bounds a flowpath for aft moving fluid. The surface has a circumferentially extending recess radially disposed relative to the airfoils. The recess has a generally aft facing wall, a generally axially directed wall, and a generally forward facing wall. The aft facing wall is oriented so as to provide a barrier to the forward flow of fluid in the clearance between airfoil and surface. The forward facing wall is oriented so as to provide an aerodynamically smooth transition from the recess into the flowpath.

Cooling Airflow Studies at the Leading Edge of a Film-Cooled Airfoil

Abstract: An experimental flow study on cooling holes in cylindrical models simulating the leading edge of a typical turbine airfoil is presented. The effect of external flow around the cylinder on the coolant discharge through a single hole is represented as a function of the momentum ratio of the cooling jet to the local external flow. A similar correlation was found for the effect of internal axial flow. The ability to separate the entrance and exit effects on the hole is due to the fact that the hole is a long orifice. The entrance and exit effects on the coolant flow are expressed as loss coefficients analogous to traditional loss coefficients in pipe flow. The loss coefficients for single holes were used to predict the total and individual flows through an array of holes in the presence of an external flow field. The total flow is predicted accurately as compared to the results of tests on arrays of holes. It can be concluded that the interaction between adjacent cooling holes is slight. The physical model can be used for coolant optimization studies.

Extended parametric representation of compressor fans and turbines. Volume 3: MODFAN user's manual (parametric modulating flow fan)

Abstract: A modeling technique for single stage flow modulating fans or centrifugal compressors has been developed which will enable the user to obtain consistent and rapid off-design performnce from design point input. The fan flow modulation may be obtained by either a VIGV (variable inlet guide vane) or a VPF (variable pitch rotor) option. Only the VIGV option is available for the centrifugal compressor. The modeling technique has been incorporated into a time-sharing program to facilitate its use. Because this report contains a description of the input output data, values of typical inputs, and examples cases, it is suitable as a user's manual. This report is the last of a three volume set describing the parametric representation of compressor fans, and turbines. The titles of the three volumes are as follows: (1) Volume 1 CMGEN USER's Manual (Parametric Compressor Generator); (2) Volume 2 PART USER's Manual (Parametric Turbine); (3) Volume 3 MODFAN USER's Manual (Parametric Modulating Flow Fan).

Steam turbine plant

Abstract: PURPOSE:To prevent an increase in the exhaust loss at partial load by providing a steam reducing valve for each stage at the inlet of a low pressure turbine of a turbine plant, and connecting the same to a steam reducing valve at the inlet of a high pressure turbine. CONSTITUTION:Steam reducing valves 8-11 are provided at the inlet of each stage of a low pressure turbine of a turbine plant having a high pressure turbine 2, intermediate pressure turbine 3 and low pressure turbine 4-7. When the load is lowered and the exhaust loss is increased, the valve 11 is almost fully closed at a set point and only a minimum flow is directed through the low pressure turbine 7. As such, the pressure of each stage of the low pressure turbine is increased to increase the velocity of the axial flow of the exhaust at the outlet of the final stage, resulting in a reduction of the exhaust loss. Also, when the valve 11 is closed, a signal is sent to a valve 1 so that the valve 1 may be opened in an interlocking manner to reduce a throttle loss.