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Showing papers on "Axial compressor published in 1995"


Journal ArticleDOI
TL;DR: In this article, an elastic plate is chosen to model the mechanics of the soft palate during oronasal snoring and the stability of the plate is investigated through an initial value problem.

274 citations


Book
01 Nov 1995
TL;DR: In this article, the authors present a detailed discussion of the properties of compressors and their properties, as well as a comparison between conventional and non-linear compressors, with respect to their performance.
Abstract: PREFACE. ACKNOWLEDGMENTS. PART I POSITIVE DISPLACEMENT COMPRESSOR TECHNOLOGY. 1 Theory. 1.1 Symbols. 1.2 How a Compressor Works. 1.3 First Law of Thermodynamics. 1.4 Second Law of Thermodynamics. 1.5 Ideal or Perfect Gas Laws. 11.6 Vapor Pressure. 1.7 Gas and Vapor. 1.8 Partial Pressures. 1.9 Critical Conditions. 1.10 Compressibility. 1.11 Generalized Compressibility Charts. 1.12 Gas Mixtures. 1.13 The Mole. 1.14 Specific Volume and Density. 1.15 Volume Percent of Constituents. 1.16 Molecular Weight of a Mixture. 1.17 Specific Gravity and Partial Pressure. 1.18 Ratio of Specific Heats. 1.19 Pseudo-critical Conditions and Compressibility. 1.20 Weight-Basis Items. 1.21 Compression Cycles. 1.22 Power Requirement. 1.23 Compressibility Correction. 1.24 Multiple Staging. 1.25 Volume References. 1.26 Cylinder Clearance and Volumetric Efficiency. 1.27 Cylinder Clearance and Compression Efficiency. Reference. 2 Reciprocating Process Compressor Design Overview. 2.1 Crankshaft Design. 2.2 Bearings and Lubrication Systems. 2.3 Connecting Rods. 2.4 Crossheads. 2.5 Frames and Cylinders. 2.6 Cooling Provisions. 2.7 Pistons. 2.8 Piston and Rider Rings. 2.9 Valves. 2.10 Piston Rods. 2.11 Packings. 2.12 Cylinder Lubrication. 2.13 Distance Pieces. 2.14 Reciprocating Compressor Modernization. 3 Reciprocating Compressor Performance and Monitoring Considerations. 3.1 Capacity Control. 3.2 More About Cylinder Jacket Cooling and Heating Arrangements. 3.3 Comparing Lubricated and Nonlubricated Conventional Cylinder Construction. 3.4 Compressor Vent and Buffer Systems. 3.5 Compressor Instrumentation. 3.6 Condition Monitoring of Reciprocating Compressors. References. 4 Labyrinth Piston Compressors. 4.1 Main Design Features. 4.2 Energy Consumption. 4.3 Sealing Problems. 5 Hypercompressors. 5.1 Introduction. 5.2 Cylinders and Piston Seals. 5.3 Cylinder Heads and Valves. 5.4 Drive Mechanism. 5.5 Miscellaneous Problems. 5.6 Conclusions. 6 Metal Diaphragm Compressors. 6.1 Introduction. 6.2 Terminology. 6.3 Description. 7 Lobe and Sliding Vane Compressors. 8 Liquid Ring Compressors. 9 Rotary Screw Compressors and Filter Separators. 9.1 Twin-Screw Machines. 9.2 Oil-Flooded Single-Screw Compressors. 9.3 Selecting Modern Reverse-Flow Filter-Separator Technology. 10 Reciprocating Compressor Performance and Sizing Fundamentals. 10.1 Theoretical Maximum Capacity. 10.2 Capacity Losses. 10.3 Valve Preload. 10.4 Valve and Gas Passage Throttling. 10.5 Piston Ring Leakage. 10.6 Packing Leakage. 10.7 Discharge Valve Leakage. 10.8 Suction Valve Leakage. 10.9 Heating Effects. 10.10 Pulsation Effects. 10.11 Horsepower. 10.12 Horsepower Adders. 10.13 Gas Properties. 10.14 Alternative Equations of State. 10.15 Condensation. 10.16 Frame Loads. 10.17 Compressor Displacement and Clearance. 10.18 Staging. 10.19 Fundamentals of Sizing. 10.20 Sizing Examples. PART II DYNAMIC COMPRESSOR TECHNOLOGY. 11 Simplified Equations for Determining the Performance of Dynamic Compressors. 11.1 Nonoverloading Characteristics of Centrifugal Compressors. 11.2 Stability. 11.3 Speed Change. 11.4 Compressor Drive. 11.5 Calculations. 12 Design Considerations and Manufacturing Techniques. 12.1 Axially vs. Radially Split. 12.2 Tightness. 12.3 Material Stress. 12.4 Nozzle Location and Maintenance. 12.5 Design Overview. 12.6 Bearing Configurations. 12.7 Casing Design Criteria. 12.8 Casing Manufacturing Techniques. 12.9 Stage Design Considerations. 12.10 Impeller Manufacturing Techniques. 12.11 Rotor Dynamic Considerations. 12.12 Fouling Considerations and Coatings. 13 Advanced Sealing and Bearing Systems. 13.1 Background. 13.2 Dry Seals. 13.3 Magnetic Bearings. 13.4 Development Efforts. 13.4.1 Thrust-Reducing Seals. 13.5 Integrated Designs. 13.6 Fluid-Induced Instability and Externally Pressurized Bearings. References. Suggested Reading. 14 Couplings, Torque Transmission, and Torque Sensing. 14.1 Coupling Overview. 14.2 Coupling Retrofits and Upgrades. 14.3 Performance Optimization Through Torque Monitoring. 15 Lubrication, Sealing, and Control Oil Systems for Turbomachinery. 15.1 Considerations Common to All Systems. 15.2 Seal Oil Considerations. 16 Compressor Control. 16.1 Introduction. 16.2 Control System Objectives. 16.3 Compressor Maps. 16.4 Performance Control. 16.5 Performance Limitations. 16.6 Preventing Surge. 16.7 Loop Decoupling. 16.8 Conclusions. Reference. 17 Head-Flow Curve Shape of Centrifugal Compressors. 17.1 Compressor Stage. 17.2 Elements of the Characteristic Shape. 17.3 Conclusions. 18 Use of Multiple-Inlet Compressors. 18.1 Critical Selection Criteria. 18.2 Design of a Sideload Compressor. 18.3 Testing. 19 Compressor Performance Testing. 19.1 Performance Testing of New Compressors. 19.2 Shop Testing and Types of Tests. 19.3 Field Testing. 19.4 Predicting Compressor Performance at Other Than As-Designed Conditions. References. 20 Procurement, Audit, and Asset Management Decisions. 20.1 Incentives to Buy from Knowledgeable and Cooperative Compressor Vendors. 20.2 Industry Standards and Their Purpose. 20.3 Disadvantages of Cheap Process Compressors. 20.4 Audits vs. Reviews. 20.5 Auditing and Reviewing Compressors. 20.6 Compressor Inspection: Extension of the Audit Effort. 20.7 Compressor Installation Specifications. Special-Purpose Machinery. References. 21 Reliability-Driven Asset Management Strategies. 21.1 Strategy for Reciprocating Compressors. 21.2 Achieving Compressor Asset Optimization. References. APPENDIX A PROPERTIES OF COMMON GASES. APPENDIX B SHORTCUT CALCULATIONS AND GRAPHICAL COMPRESSOR SELECTION PROCEDURES. APPENDIX C BIBLIOGRAPHY AND LIST OF CONTRIBUTORS. INDEX.

138 citations


Patent
02 Nov 1995
TL;DR: In this article, a technique for controlling compressor stall and surge is disclosed, where static pressure asymmetry is sensed at a plurality of locations along the circumference of the compressor inlet.
Abstract: A technique for controlling compressor stall and surge is disclosed. In a gas turbine engine, static pressure asymmetry is sensed at a plurality of locations along the circumference of the compressor inlet. Time rate of change of the mass flow in the compressor is also estimated using pressure measurements in the compressor. A signal processor uses these signals to modulate a compressor bleed valve responsive to the level of flow property asymmetry, the time rate of change of the annulus average flow to enhance operability of the compressor.

127 citations


Journal ArticleDOI
TL;DR: In this article, the performance degradation of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported, where a 0.025 mm (0.001 in) thick rough coating with a surface finish of 2.54-3.18 rms (100-125 rms μm) is applied to the rotor blades.
Abstract: The performance deterioration of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54-3.18 rms μm (100-125 rms μin.) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10 percent at the hub and 20 percent at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9 percent reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254-0.508 rms μm (10-20 rms μin.), compared to the bare metal blade surface finish of 0.508 rms μm (20 rms μin.). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60, 80, and 100 percent of design speed. The results indicate that thickness/roughness over the first 2 percent of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70 percent of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-three-dimensional Navier-Stokes flow solver, which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that adding roughness at the blade leading edge causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage, which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

119 citations


Journal ArticleDOI
TL;DR: In this article, the axial, tangential, and radial components of relative velocity, as well as the static and stagnation pressures, were obtained at two axial locations, one at the rotor trailing edge, the other downstream of the rotor.
Abstract: Detailed measurements of the flow field in the tip region of an axial flow compressor rotor were carried out using a rotating five-hole probe. The axial, tangential, and radial components of relative velocity, as well as the static and stagnation pressures, were obtained at two axial locations, one at the rotor trailing edge, the other downstream of the rotor. The measurements were taken up to about 26 percent of the blade span from the blade tip. The data are interpreted to understand the complex nature of the flow in the tip region, which involves the interaction of the tip leakage flow, the annulus wall boundary layer and the blade wake. The experimental data show that the leakage jet does not roll up into a vortex. The leakage jet exiting from the tip gap is of high velocity and mixes quickly with the mainstream, producing intense shearing and flow separation. There are substantial differences in the structure of tip clearance observed in cascades and rotors.

115 citations


Patent
31 Mar 1995
TL;DR: In this paper, a gas turbine including a compressor having a bore and a rotor comprised of multiple stages extending between a first stage at a forward end of the compressor and a last stage at an aft end of a compressor, each stage including a rotor disk having a peripheral rim and multiple blades secured to the peripheral rim, a combustion system comprising a plurality of combustors utilizing discharge air from the compressor for combustion, and multiple turbine stages driven by combustion gases from the combustion system, is described.
Abstract: In a gas turbine including a compressor having a bore and a rotor comprised of multiple stages extending between a first stage at a forward end of the compressor and a last stage at an aft end of the compressor, each stage including a rotor disk having a peripheral rim and multiple blades secured to the peripheral rim, a combustion system comprising a plurality of combustors utilizing discharge air from the compressor for combustion, and multiple turbine stages driven by combustion gases from the combustion system, the improvement comprising means for supplying cooling air at least to a peripheral rim of the last stage of the compressor.

108 citations


Patent
29 Aug 1995
TL;DR: In this paper, a multiple stage supercharging system for both two-stroke and four-stroke internal combustion engines is described, which is suitable for both four and five-stroke engines.
Abstract: A multiple stage supercharging system is disclosed suitable for both two-stroke cycle and four-stroke cycle internal combustion engines Ambient air, which is propelled by the forward velocity of the engine, enters an air cleaner housing (11) through an air filter The air cleaner housing (11) attaches to the air intake (32) of a centrifugal compressor (12) The centrifugal compressor (12) mounts directly to the magnetic flywheel on the crankshaft of the engine The centrifugal compressor wheel (22) pressurizes the ambient air for use in the combustion process The outlet of the centrifugal compressor housing mates with a secondary plenum chamber (17) The outlet of the secondary plenum chamber (17) mates with a dc motor driven axial compressor (28) The axial compressor (28) operates on current derived from a motor driven alternator (38) The outlet of the axial compressor connects to a primary plenum chamber (18) which connects to the air intake snorkel on the carburetor A pressure equalization tube (19) extends from the primary plenum chamber to the carburetor bowl to allow for consistent flow of the air/fuel mixture to the crankcase The system provides for multiple compressors to generate layers of additive pressure for supercharging an internal combustion engine The system provides air pressure to boost the power output of the engine across the entire rpm band by utilizing the centrifugal compressor (12) and the axial compressor (28) at low speeds and by utilizing forward air velocity air intake pressure plus the centrifugal and axial compressors at high speeds

97 citations


Patent
23 Oct 1995
TL;DR: A gas turbine engine includes a compressor, combustor, and turbine, and a cooling air closed-circuit having an extraction line from the compressor to a hot turbine component such as vanes, blades, or shrouds as discussed by the authors.
Abstract: A gas turbine engine includes a compressor, combustor, and turbine, and a cooling air closed-circuit having an extraction line from the compressor to a hot turbine component such as vanes, blades, or shrouds, and a return line to the compressor The extraction line may be joined to any suitable compressor stage for extracting air at a suitable pressure for flow to the turbine, with the return line being joined to any suitable injection stage of the compressor having a lower pressure than the extraction stage for driving the cooling air through the closed-circuit The compressor itself therefore drives the closed-circuit and is self regulated

83 citations


Proceedings ArticleDOI
05 Jun 1995
TL;DR: In this article, the boundary layer characteristics of axial flow compressors and LP turbines were analyzed using hot wire probes. But the results were limited to a single-stage compressor and turbine.
Abstract: Comprehensive experiments and computational analyses were conducted to understand boundary layer development on airfoil surfaces in multistage, axial-flow compressors and LP turbines. The tests were run over a broad range of Reynolds numbers and loading levels in large, low-speed research facilities which simulate the relevant aerodynamic features of modern engine components. Measurements of boundary layer characteristics were obtained by using arrays of densely packed, hot-film gauges mounted on airfoil surfaces and by making boundary layer surveys with hot wire probes. Computational predictions were made using both steady flow codes and an unsteady flow code. This is the first time that time-resolved boundary layer measurements and detailed comparisons of measured data with predictions of boundary layer codes have been reported for multistage compressor and turbine blading.Part 1 of this paper draws a composite picture of boundary layer development in turbomachinery based upon a synthesis of all of our experimental findings for the compressor and turbine. Parts 2 and 3 present the experimental results for the compressor and turbine, respectively. Part 4 presents computational analyses and discusses comparisons with experimental data.For both compressor and turbine blading, the experimental results show large extents of laminar and transitional flow on the suction surface of embedded stages, with the boundary layer generally developing along two distinct but coupled paths. One path lies approximately under the wake trajectory while the other lies between wakes. Along both paths the boundary layer clearly goes from laminar to transitional to turbulent. The wake path and the non-wake path are coupled by a calmed region which, being generated by turbulent spots produced in the wake path, is effective in suppressing flow separation and delaying transition in the non-wake path. The location and strength of the various regions within the paths, such as wake-induced transitional and turbulent strips, vary with Reynolds number, loading level and turbulence intensity. On the pressure surface, transition takes place near the leading edge for the blading tested. For both surfaces, bypass transition and separated-flow transition were observed. Classical Tollmien-Schlichting transition did not play a significant role. Comparisons of embedded and first-stage results were also made to assess the relevance of applying single-stage and cascade studies to the multistage environment.Although doing well under certain conditions, the codes in general could not adequately predict the onset and extent of transition in regions affected by calming. However, assessments are made to guide designers in using current predictive schemes to compute boundary layer features and obtain reasonable loss predictions.Copyright © 1995 by ASME

82 citations


Proceedings ArticleDOI
30 May 1995
TL;DR: In this article, a nonlinear control strategy is designed based on an analytical model to achieve simultaneous active control of rotating stall and surge in an axial flow compression system with relevant dynamics representative of modern aeroengines.
Abstract: Aeroengines operate in regimes for which both rotating stall and surge impose low flow operability limits. Thus, active control strategies designed to enhance operability of aeroengines must address both rotating stall and surge as well as their interaction. In this paper, a nonlinear control strategy is designed based on an analytical model to achieve simultaneous active control of rotating stall and surge in an axial flow compression system with relevant dynamics representative of modern aeroengines. The controller is experimentally validated on a 3-stage low-speed axial flow compression system. This rig is dynamically scaled to replicate the interaction between rotating stall and surge typical of modern aeroengines, and several experimental results are presented for this rig. For actuation, the control stategy utilizes a single plenum bleed valve with bandwidth on the order of the rotor frequency. For sensing, measurements of the circumferential asymmetry and annulus-averaged unsteadiness of the flow through the compressor are used. Experimental validation of simultaneous control of rotating stall and surge with minimal sensing and actuation requirements is viewed as an important step towards applying active control to enhance operability of compression systems in modern aeroengines.

78 citations


Journal ArticleDOI
TL;DR: In this article, the Squire-Long equation is used to investigate the dependence of solutions on upstream, or inlet, and downstream, or outlet, boundary conditions and flow geometry.
Abstract: The steady axisymmetric Euler flow of an inviscid incompressible swirling fluid is described exactly by the Squire-Long equation. This equation is studied numerically for the case of diverging flow to investigate the dependence of solutions on upstream, or inlet, and downstream, or outlet, boundary conditions and flow geometry. The work is performed with a view to understanding how the phenomenon of vortex breakdown occurs. It is shown that solutions fail to exist or, alternatively, that the axial flow ceases to be unidirectional, so that breakdown can be inferred, when a parameter measuring the relative magnitude of rotation and axial flow (the Squire number) exceeds critical values depending upon the geometry and inlet profiles. A 'quasi-cylindrical' simplification of the Squire-Long equation is compared with the more complete Euler model and shown to be able to account for most of the latter's behaviour. The relationship is examined between 'failure' of the quasi-cylindrical model and the occurrence of a 'critical' flow state in which disturbances can stand in the flow.

Journal Article
TL;DR: In this article, the performance of an impulse turbine with self-pitch-controlled guide vanes for wave power conversion has been investigated, and the results show that a high-efficiency impulse turbine can be developed by these of guide vane connected by link motions.
Abstract: Experimental investigations directed towards improving the performance of the impulse turbine with self-pitch-controlled guide vanes for wave power conversion are reported. The turbine presented and tested here has an upstream and a downstream guide vane row connected by link motions. The behavior of guide vanes in the reciprocating flows is shown in connection with the axial flow velocity. The results show that a high-efficiency impulse turbine can be developed by these of guide vanes connected by link motions. Furthermore, it is found that the running and starting characteristics of this turbine in the reciprocating flow can be estimated from the performance of the turbine with fixed nozzle and diffuser vanes in a unidirectional steady flow.

Journal ArticleDOI
C. Y. Wu1
TL;DR: In this article, it was shown that many complex arbitrary surfaces typical of our blades in fans, axial compressors, and centrifugal impellers in aviation gas turbines can be rendered exactly flank millable with one or more passes per surface often without sacrificing, indeed with gain, in performance.
Abstract: It is generally conceived that a blade surface is flank millable if it can be closely approximated by a ruled surface ; otherwise the slow machining process of point milling has to be employed. However, we have now demonstrated that the ruled surface criterion for flank milling is neither necessary nor sufficient. Furthermore, many complex arbitrary surfaces typical of our blades in fans, axial compressors, and centrifugal impellers in aviation gas turbines are actually closely flank millable and can be rendered exactly flank millable with one or more passes per surface often without sacrificing, indeed usually with gain, in performance.

Proceedings ArticleDOI
13 Dec 1995
TL;DR: In this article, the effect of a compressor characteristic actuation scheme for the three-state Moore Greitzer compression system model was analyzed and the closed loop feedback based on the square magnitude of the first rotating stall mode was used to decrease the hysteresis region associated with the transition from unstalled to stalled and back to unstalled operation.
Abstract: Previous results in the use of pulsed air injection for active control of rotating stall have suggested that air injectors have the effect of shifting the steady state compressor characteristic. In this paper we analyze the effect of a compressor characteristic actuation scheme for the three state Moore Greitzer compression system model. It is shown that closed loop feedback based on the square magnitude of the first rotating stall mode can be used to decrease the hysteresis region associated with the transition from unstalled to stalled and back to unstalled operation. The compressor characteristic shifting idea is then applied to a higher fidelity distributed model in which the characteristic shifting has phase content in addition to the magnitude content captured by the three state model. The optimal phasing of the air injection relative to the sensed position of the stall cell is determined via simulation and the results found to agree with those obtained via an experimental parametric study on the Caltech low-speed axial flow compressor.

Journal ArticleDOI
TL;DR: In this article, a simplified expression is derived for the separation efficiency, defined as the maximum length needed to remove droplets or particles form the flow, and a design in which the swirl generator is mounted on a central body positioned in a wider section of the pipe than the separation section is presented.
Abstract: The design of an axial cyclone for the removal of liquid droplets from a liquid stream is considered. The design constraints are that the swirl in the separation section is to be maximized and that an adverse pressure gradient in the flow is to be avoided everywhere in order to prevent droplet breakup. First, a simplified expression is derived for the separation efficiency, defined as the maximum length needed to remove droplets or particles form the flow. The result shows that it is advantageous to position the vanes that generate the swirl at a radius larger than the section where actual separation takes place. This leads to a design in which the swirl generator is mounted on a central body positioned in a wider section of the pipe than the separation section. To calculate the flow and the pressure field in this cyclone geometry, the authors develop a stream function method which uses a superposition of axisymmetric sources and sinks. They show by means of an example that they can indeed design a geometry in which the undesirable adverse pressure gradients is avoided.

Journal ArticleDOI
TL;DR: In this paper, a one-dimensional unsteady compressible viscous flow model of a generic compression system was applied to a multistage axial compressor experimental rig configured for single-stage operation.
Abstract: A one-dimensional unsteady compressible viscous flow model of a generic compression system previously developed by the authors is applied to a multistage axial compressor experimental rig configured for single-stage operation. The required model parameters and maps are identified from experimental data. The resulting model is an explicit system of nine first-order ODEs. The model inputs are compressor speed, nozzle area, compressor discharge bleed area, plenum bleed area, inlet total pressure and entropy, and nozzle and bleed exit static pressures. The model and experimental data are compared with respect to both open-loop uncontrolled and closed-loop controlled behaviors. These comparisons focus on (i) forced transients and (ii) global nonlinear dynamics and bifurcations. In all cases the agreement between the model and experimental data is excellent. Of particular interest is the ability of the model, which does not include any hysteretic maps, to predict experimentally observed hysteresis with respect to the onset and cessation of surge. This predictive capability of the model manifests itself as the coexistence of a stable equilibrium (rotating stall) and a stable periodic solution (surge) in the model at a single fixed set of system input values. Also of interest is the fact that the controllers used for closed-loop comparisons were designed directly from the model with no a posteriori tuning of controller parameters. Thus, the excellent closed-loop comparisons between the model and experimental data provide strong evidence in support of the validity of the model for use in direct model based controller design. The excellent agreement between the model and experimental data summarized above is attributed in large part to the use of effective lengths within the model, as functions of axial Mach number and nondimensional compressor rotational speed, as prescribed by the modeling technique. The use of these effective lengths proved to be far superior to the use of physical lengths. The use of these effective lengths also provided substantial improvement over the use of physical lengths coupled with fixed first-order empirical lags, as proposed by other authors for the modeling of observed compressor dynamic lag. The overall success of this model is believed to represent a positive first step toward a complete experimental validation of the approach to control-oriented high-frequency turbomachinery modeling being developed by the authors

Proceedings ArticleDOI
05 Jun 1995
TL;DR: In this article, the experimental evidence used to construct the composite picture for LP turbines that was given in the discussion in Section 5.0 of Part 1.0 is presented and interpret the data from the surface hot-film gauges and the boundary layer surveys for the baseline operating condition.
Abstract: This is Part Three of a four-part paper. It begins with Section 11.0 and continues to describe the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery.In this part, we present the experimental evidence that we used to construct the composite picture for LP turbines that was given in the discussion in Section 5.0 of Part 1. We present and interpret the data from the surface hot-film gauges and the boundary layer surveys for the baseline operating condition. We then show how this picture changes with variations in Reynolds number, airfoil loading and nozzle-nozzle clocking.Copyright © 1995 by ASME

Proceedings ArticleDOI
05 Jun 1995
TL;DR: In this article, two versions of a three dimensional multistage Navier-Stokes code were used to optimize the design of an eleven stage high pressure compressor, which contained bowed stators and rotor airfoils with contoured endwalls.
Abstract: Two versions of a three dimensional multistage Navier-Stokes code were used to optimize the design of an eleven stage high pressure compressor. The first version of the code utilized a “mixing plane” approach to compute the flow through multistage machines. The effects due to tip clearances and flowpath cavities were not modeled. This code was used to minimize the regions of separation on airfoil and endwall surfaces for the compressor. The resulting compressor contained bowed stators and rotor airfoils with contoured endwalls. Experimental data acquired for the HPC showed that it achieved 2% higher efficiency than a baseline machine, but it had 14% lower stall margin. Increased stall margin of the HPC was achieved by modifying the stator airfoils without compromising the gain in efficiency as demonstrated in subsequent rig and engine tests. The modifications to the stators were defined by using the second version of the multistage Navier-Stokes code, which models the effects of tip clearance and endwall flowpath cavities, as well as the effects of adjacent airfoil rows through the use of “bodyforces” and “deterministic stresses”. The application of the Navier-Stokes code was assessed to yield up to 50% reduction in the compressor development time and cost.© 1995 ASME


Journal ArticleDOI
TL;DR: In this article, the aeromechanical feedback was used to extend the stable operating range of an axial flow compressor, and the first use of local feedback and dynamic compensation techniques to suppress rotating stall.
Abstract: Dynamic control of rotating stall in an axial flow compressor has been implemented using aeromechanical feedback. The control strategy developed used an array of wall jets, upstream of a single-stage compressor, which were regulated by locally reacting reed valves. These reed valves responded to the small-amplitude flow-field pressure perturbations that precede rotating stall. The valve design was such that the combined system, compressor plus reed valve controller, was stable under operating conditions that had been unstable without feedback. A 10 percent decrease in the stalling flow coefficient was obtained using the control strategy, and the extension of stable flow range was achieved with no measurable change in the steady-state performance of the compression system. The experiments demonstrate the first use of aeromechanical feedback to extend the stable operating range of an axial flow compressor, and the first use of local feedback and dynamic compensation techniques to suppress rotating stall. The design of the experiment was based on a two-dimensional stall inception model, which incorporated the effect of the aeromechanical feedback. The physical mechanism for rotating stall in axial flow compressors was examined with focus on the role of dynamic feedback in stabilizing compression system instability. As predicted and experimentally demonstrated, the effectiveness of the aeromechanical control strategy depends on a set of nondimensional control parameters that determine the interaction of the control strategy and the rotating stall dynamics.

PatentDOI
TL;DR: In this paper, the axial fan for a cooling blower of a vehicle engine has fins with leading and trailing edges, and aeroacoustic optimization is provided by each of the leading edges and the trailing edges having a strong forward sweep followed by a strong backward sweep in the manner of a bird's wing.
Abstract: An axial fan for a cooling blower of a vehicle engine has fins with leading and trailing edges, and aeroacoustic optimization is provided by each of the leading edges and trailing edges having a strong forward sweep followed by a strong backward sweep in the manner of a bird's wing or by a straight forward sweeps followed by a strong backward sweep.

Journal ArticleDOI
TL;DR: In this article, the aerodynamics of a rotorcraft in low-speed flight with the airframe must be modeled, and various phenomena encountered during such interactions are summarized, combining previous results on various configurations with recent experimental results.
Abstract: To compute the aerodynamics of a rotorcraft in low-speed flight, the interaction of the strong vortices in the rotor wake with the airframe must be modeled. Using a hemisphere-cylinder airframe and a two-bladed rotor for reference, the various phenomena encountered during such interactions are summarized, combining previous results on various configurations with recent experimental results. Differences between the interaction at the front and aft portions of the wake are discussed. The precollision phase conforms to expectations from potential flow, and includes distortion of the vortex trajectory determined by the sense of rotation of the vortex. The collision phase involves complex boundary-layer interactions. The axial velocity in the vortex core causes substantial asymmetry and influences the surface pressure distribution on the airframe side under the advancing rotor blade, where the axial flow stagnates. The postinteraction vortex is much weaker, but still contains some swirl energy. Where flow separation occurs due to airframe shapes, the interaction is not modified significantly, because the vortex dominates the interaction with separated shear layers for parameter values of practical interest. Areas of remaining uncertainty are discussed.

Journal ArticleDOI
TL;DR: In this article, an apparatus designed to subject a liquid column or bridge to an axial flow is described, where the bridge liquid is density matched to the water in the external flow in order to simulate low gravity (∼10 -3 Earth g).

Journal ArticleDOI
TL;DR: In this paper, a high-Reynolds number pump (HIREP) was used to acquire flow measurements in the rotor blade tip clearance region, with blade chord Reynolds numbers of 3,900,000 and 5,500,000.
Abstract: A high-Reynolds-number pump (HIREP) facility has been used to acquire flow measurements in the rotor blade tip clearance region, with blade chord Reynolds numbers of 3,900,000 and 5,500,000. The initial experiment involved rotor blades with varying tip clearances, while a second experiment involved a more detailed investigation of a rotor blade row with a single tip clearance. The flow visualization on the blade surface and within the flow field indicate the existence of a trailing-edge separation vortex, a vortex that migrates radially upward along the trailing edge and then turns in the circumferential direction near the casing, moving in the opposite direction of blade rotation. Flow visualization also helps in establishing the trajectory of the tip leakage vortex core and shows the unsteadiness of the vortex. Detailed measurements show the effects of tip clearance size and downstream distance on the structure of the rotor tip leakage vortex. The character of the velocity profile along the vortex core changes from a jetlike profile to a wakelike profile as the tip clearance becomes smaller. Also, for small clearances, the presence and proximity of the casing endwall affects the roll-up, shape, dissipation, and unsteadiness of the tip leakage vortex. Measurements also show how much circulation is retained by the blade tip and how much is shed into the vortex, a vortex associated with high losses.

Patent
04 Jul 1995
TL;DR: An auxiliary gas turbine engine comprises: a load compressor 110, a cabin air supply duct 115 which supplies at least some of the air from the exit of the load compressor to the aircraft cabin; a high pressure (HP) compressor 120 which further compresses the remaining air from an exit of a load-compressor 110; a combustor 130, a turbine 140, an electrical generator 150 driven by the turbine 140; an electric motor 160 driven by electrical output of the generator 150, the electric motor160 in turn driving the load compressor 110; and at least one air heater/cool
Abstract: An auxiliary gas turbine engine comprises: a load compressor 110; a cabin air supply duct 115 which supplies at least some of the air from the exit of the load compressor 110 to the aircraft cabin; a high pressure (HP) compressor 120 which further compresses the remaining air from the exit of the load compressor 110; a combustor 130; a turbine 140; an electrical generator 150 driven by the turbine 140; an electric motor 160 driven by the electrical output of the generator 150, the electric motor 160 in turn driving the load compressor 110; and at least one air heater/cooler 170A, 170B or 170C comprising a heat exchanger matrix 172A, 172B or 172C located in the appropriate air duct and a reversible refrigeration plant or heat pump 174A, 174B or 174C which is electrically powered from the generator 150 so that it can either cool or heat the air passing through the matrix. The load compressor and heat exchanger(s) together produce air at the required pressure and temperature for supply to the aircraft cabin.

Patent
01 Aug 1995
TL;DR: In this article, a compressor stator vane assemblage includes a first metal bushing disposed within a bore through the compressor casing and bolted to the casing by externally accessible bolts.
Abstract: A compressor stator vane assemblage includes a first metal bushing disposed within a bore through the compressor casing and bolted to the casing by externally accessible bolts. A second composite bushing is disposed within the first bushing and receives the spindle of the stator vane. Reduced diameter portions of the spindle project through openings in the first and second bushings. A lever attaches to the spindle portion and is movable to rotate the vane. By removing the bolts, the first and second bushings can be removed from the casing for replacement or rotation of 180° for prolonged service life without disassembly of the casing or removal of the stator vane.

Patent
Ronald L. Haugen1
30 May 1995
TL;DR: In this article, a check valve is connected with the discharge for preventing high pressure fluid from back flowing to the compressor, and a vane diffuser assembly fluidly communicates with the impeller.
Abstract: An apparatus achieves passive damping of flow disturbances to control centrifugal compressor surge. The apparatus includes a centrifugal compressor for compressing a low pressure fluid. The centrifugal compressor has an impeller, an inlet which communicates with an atmosphere and a discharge through which compressed air is supplied to a compressed air system. A fluid flow control is flow connected with the inlet for controlling the flow of a low pressure fluid to the compressor. A check valve is flow connected with the discharge for preventing high pressure fluid from back flowing to the compressor. A vane diffuser assembly fluidly communicates with the impeller. A spring-mass-damper system is coupled to any one or all of the fluid flow control, check valve or vane diffuser to dampen low amplitude flow disturbances of the compressible fluid.

Journal ArticleDOI
TL;DR: A fluid dynamics analysis using flow visualization was performed to investigate the flow fields and to determine areas within the pump that could be improved, showing that the flow straightener exacerbates a discontinuity found between it and the impeller.
Abstract: The Baylor/NASA Axial Blood Flow Pump has been developed for use as an implantable left ventricular assist device (LVAD). The pump is intended as an assist device for either pulmonary or systemic circulatory support for more than 3-months' duration. To date the pump provides acceptable results in terms of thrombus formation and hemolysis (IH of 0.018 g/100 L). A fluid dynamics analysis using flow visualization was performed to investigate the flow fields and to determine areas within the pump that could be improved. These studies focused upon the inflow area in front of the pump. A prototype axial flow pump assembly was constructed to facilitate the flow visualization studies. Particle image tracking velocimetry techniques were used to measure Amberlite particles suspended in a blood analog fluid composed of 63% water and 37% glycerin. This method used a pulsed (612 Hz) laser light to determine flow velocity profiles, shear stress, Reynolds numbers, and stagnant areas within the axial pump. These studies showed that the flow straightener (a vaned assembly in the pump inflow) reduced Reynolds numbers from 4,640 to 2,540 (at 8.5 L/min) and that the flow straightener exacerbates a discontinuity found between it and the impeller. Within the inflow area, a maximum of 80 N/m2 shear stress was measured, which is well below published blood damage thresholds. Design variations were investigated resulting in a smoother flow transition between flow straightener and impeller. These variations must be investigated further to establish a correlation with hemolysis and thrombus formation.

Patent
28 Feb 1995
TL;DR: In this paper, an integrated system for controlling the positioning of a rotating rotor within a compressor to enhance the performance of a gas turbine engine is presented. But the rotor is attached to a rotating shaft that is coupled to the compressor housing by a magnetic bearing system, and the magnetic bearings support and position the rotor within the high speed tubomachine.
Abstract: An integrated system for controlling the positioning of a rotating rotor within a compressor to enhance the performance of a gas turbine engine. The rotor being attached to a rotating shaft that is coupled to the compressor housing by a magnetic bearing system. The magnetic bearings support and position the rotor within the high speed tubomachine. During operation of the compressor the magnetic bearings non-axially position the rotor relative to the compressor housing to provide compressor instability control. By non-axially displacing the rotor relative to the compressor housing there is a resulting local disturbance to the fluid flow field, which in turn serves as the disturbance to alter the dynamics of the aerodynamic system and displace the compressor surge line.

Journal ArticleDOI
TL;DR: In this paper, the rotor exit flow field was measured with an unsteady hot-wire probe, which has high temporal and spatial resolution, and the steady calculation did not predict the secondary flow at the stage exit accurately.
Abstract: Detailed experimental and numerical studies have been performed in a subsonic, axial-flow turbine stage to investigate the secondary flow field, the aerodynamic loss generation, and the spanwise mixing under a stage environment. The experimental study includes measurements of the static pressure distribution on the rotor blade surface and the rotor exit flow field using three-dimensional hot-wire and pneumatic probes. The rotor exit flow field was measured with an unsteady hot-wire probe, which has high temporal and spatial resolution. Both steady and unsteady numerical analyses were performed with a three-dimensional Navier-Stokes code for the multiple blade rows. Special attention was focused on how well the steady multiple-blade-row calculation predicts the rotor exit flow field and how much the blade interaction affects the radial distribution of flow properties at the stage exit. Detailed comparisons between the measurement and the steady calculation indicate that the steady multiple-blade-row calculation predicts the overall time-averaged flow field very well. However, the steady calculation does not predict the secondary flow at the stage exit accurately. The current study indicates that the passage vortex near the hub of the rotor is transported toward the midspan due to the blade interaction effects. Also, the structure of the secondaryflowfield at the exit of the rotor is significantly modified by the unsteady effects. The time-averaged secondary flow field and the radial distribution of the flow properties, which are used for the design of the following stage, can be predicted more accurately with the unsteady flow calculation.