Showing papers on "Axial compressor published in 1971"
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TL;DR: In this paper, a theory for the production and propagation of the shock wave associated with the supersonic elements of the axial flow compressor is presented, showing that the strength and positions of these shock waves are sensitive to small blade-to-blade differences found in practical builds.
96 citations
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TL;DR: In this article, the authors reviewed the physical mechanisms involved in each of these processes, including the generating of unsteady lift by turbulence, and compared the sound power estimates, where possible, with experiment results.
Abstract: Experiments and theory relating to fan noise sources are reviewed with emphasis on axial flow machines At supersonic rotor speeds, the steady shock pattern attached to a rotor is an efficient radiator of sound In most practical cases of subsonic rotor operation, however, direct radiation from the rotor‐locked pressure field is negligible compared with the indirect radiation, or scattering, caused by circumferential distortions in the steady flow field surrounding the rotor Random timewise modulation of the distortion changes the scattered spectrum from discrete to continuous, with a gradual progression from narrow‐band tones to broad‐band noise as the modulation bandwidth is increased Similar scattering occurs when a non‐uniform unsteady flow impinges on stator vanes, but here the radiated frequency is that of the impinging flow Finally, for blades operating in flows free from circumferential distortions, self‐generated turbulence becomes an important source of noise The paper describes the physical mechanisms involved in each of these processes, including the generating of unsteady lift by turbulence Order‐of‐magnitude sound power estimates are compared, where possible, with experiment
91 citations
01 Nov 1971
TL;DR: In this paper, an axial flow compressor rotor was tested at design speed with six different casing treatments across the rotor tip, and radial surveys of pressure, temperature, and flow angle were taken at the rotor inlet and outlet.
Abstract: An axial flow compressor rotor was tested at design speed with six different casing treatments across the rotor tip. Radial surveys of pressure, temperature, and flow angle were taken at the rotor inlet and outlet. Surveys were taken at several weight flows for each treatment. All the casings treatments decreased the weight flow at stall over that for the solid casing. Radial surveys indicate that the performance over the entire radial span of the blade is affected by the treatment across the rotor tip.
74 citations
01 Nov 1971
TL;DR: Several geometrically different porous casings were tested with an axial-flow compressor rotor to determine their effects upon the rotor stall-limit line and overall performance as discussed by the authors, and the rotor performance with the various casing treatments is compared with that obtained with a solid casing.
Abstract: Several geometrically different porous casings were tested with an axial-flow compressor rotor to determine their effects upon the rotor stall-limit line and overall performance. The tests were conducted using both uniform and nonuniform inlet-flow conditions. The rotor performance with the various casing treatments is compared with that obtained with a solid casing. The ability of the various casing treatments to displace the rotor stall-limit line to lower weight flows was observed. Significant stall-margin increases were obtained with several of the porous casings. Peak efficiencies with two of the porous casings were as high as or slightly higher than that obtained with solid casing.
64 citations
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26 Jul 1971TL;DR: In this paper, an axial flow gas turbine has a structure including a fuel nozzle for providing any desired turbine inlet temperature profile according to the mechanical stress on the rotating turbine blades.
Abstract: An axial flow gas turbine having a structure including a fuel nozzle for providing any desired turbine inlet temperature profile according to the mechanical stress on the rotating turbine blades. The fuel nozzle may be a combined multiple fuel gas and liquid type or a single fuel type, an important feature of both being the provision of fuel injection asymmetrically into the combustion chamber to establish a desired ignited fuel temperature pattern therein which continues down to the turbine inlet. This asymmetrical fuel supply into the combustion chamber is obtained by providing a number or size of fuel jets according to the temperature desired; for example, to provide a temperature gradient decreasing in an inward turbine radial direction through the combustor to correspond to the desired turbine blade inlet temperature profile. The angular jet direction also may be varied for different parts of the fuel injection pattern to obtain a further spatial control of the fuel injection distribution into the combustor. These fuel distribution control variables may be used singly or in any combination, and the same or different ones may be used for the gas and liquid jets in multiple fuel nozzles. In addition, a further modification of the intake temperature profile may be obtained by providing relatively cool jets of gaseous fluid through spaced orifices in the radially inner sides of the transition passages which direct the high temperature working gas to the turbine blades.
51 citations
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TL;DR: In this paper, a two-stage transonic research compressor has been investigated experimentally over a range of tip relative Mach numbers up to 1.56 and the results show that the phenomenon is due to the propagation at supersonic relative tip speeds of the steady rotating pressure field associated with the first-stage rotor blades.
Abstract: The buzz-saw noise made by a two-stage transonic research compressor has been investigated experimentally over a range of tip relative Mach numbers up to 1.56. The results show that the phenomenon is due to the propagation at supersonic relative tip speeds of the steady rotating pressure field associated with the first-stage rotor blades. The flow entering the tip section of the rotor has been analyzed theoretically and the circumferential pressure fluctuations computed, with good agreement with near-field measurements. The analysis leads to a clearer understanding of the dependence of the noise on inlet Mach number and three-dimensional effects and indicates the types of rotor irregularity which will most influence the harmonic content.
49 citations
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04 Jan 1971
TL;DR: In this article, a multiple piston gas compressor for use in air conditioning systems for vehicles in which the inlet of the compressor receives a mixture of refrigerant gas and oil in suspension that is impinged upon a novel rotor or swash plate is described.
Abstract: A multiple piston refrigerant gas compressor for use in air conditioning systems for vehicles in which the inlet of the compressor receives a mixture of refrigerant gas and oil in suspension that is impinged upon a novel rotor or swash plate, whereby the oil suspended in the refrigerant is separated by the centrifugal action of the rotor or swash plate and is distributed to the respective areas of the compressor requiring lubrication. The excess oil, if any, is slung outwardly by centrifugal force against the walls of the compressor housing.
46 citations
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26 Aug 1971
TL;DR: In this article, the authors propose a method for bleding from one or several compressors in the engine for cooling engine components and providing a seal between the rotary and stationary assemblies of the several rotor systems, and in which means are additionally provided for compressor control.
Abstract: This invention relates to gas turbine engines, and, more particularly, to multishaft turbojet aero engines having a plurality of compressors and turbines, in which air is bled from one or several compressors in the engine for cooling engine components and for providing a seal between the rotary and stationary assemblies of the several rotor systems, and in which means are additionally provided for compressor control.
46 citations
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TL;DR: In this paper, the transmission of spinning acoustic modes through a throat in a long cylindrical duct is considered, and a substantial attenuation of wave energy is found for throats with reasonable area reduction.
39 citations
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TL;DR: In this article, a series of rotating vanes, stationary vanes or supporting structure of streamline struts is oriented in circumferentially leaning relation to another series of blades in the machine to effect a reduction in noise.
Abstract: In an air blower or axial compressor, such as employed in airplane power plants, ventilating systems, etc, a series of rotating vanes, stationary vanes, or supporting structure of streamline struts is oriented in circumferentially leaning relation to another series of blades in the machine so as to effect a reduction in noise
37 citations
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TL;DR: In this article, the effect of axial flow on the stability of a single vortex filament and a pair of vortices has been investigated and the stability boundary has been obtained.
Abstract: Previous results concerning the effects of axial velocity on the motion of vortex filaments are reviewed. These results suggest that a slender-body force balance between the Kutta–Joukowski lift on the vortex cross-section and the momentum flux within the curved filament will give some insight into the behaviour of the filament. These simple ideas are exploited for both a single vortex filament and a vortex pair, both containing axial flow. The stability of a straight vortex filament containing an axial flow to long wave sinusoidal displacements of its centre-line is investigated and the stability boundary obtained. The effect of axial flow on the stability of a vortex pair is explored. It is shown that to lowest order (in the ratio of vortex core radius to distance between the vortices) the effect of axial flow is to reduce the self-induced rotation of a single filament and that this effect can be considered as a change in effective core radius. To the next order, travelling waves appear in the instability, the instability mode for the vortex pair becomes non-planar but the amplification rate of the instability is not affected.
01 Dec 1971
TL;DR: In this paper, simplified equations for estimating the length and weight of major powerplant components of VTOL aircraft are presented, including fan, fan duct, compressor, combustor, turbine, structure, and accessories.
Abstract: Simplified equations are presented for estimating the length and weight of major powerplant components of VTOL aircraft. The equations were developed from correlations of lift and cruise engine data. Components involved include fan, fan duct, compressor, combustor, turbine, structure, and accessories. Comparisons of actual and calculated total engine weights are included for several representative engines.
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TL;DR: In this paper, the aerodynamic stability of a long, thin cylindrical shell with the outer surface exposed to an inviscid, helical flow of ah* is investigated.
Abstract: The aeroelastic stability of a long, thin cylindrical shell with the outer surface exposed to an inviscid, helical flow of ah* is investigated. The cylinder behavior is described by classical shell equations, whereas the aerodynamic forces are described by the linearized potential theory. The approach that is used herein examines the nature of stability of the system when the system is "slightly" perturbed from its initial equilibrium state. In this paper, numerical results are presented only for the special case of swirl flow around a nonrotating shell, i.e., the axial flow velocity is set to zero. These results indicate that traveling wave type of flutter can be caused by coalescence of backward and forward traveling waves. Two approximate theories are presented and the results are compared.
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19 Mar 1971TL;DR: A support, actuation and balancing structure for rotor blades in a variable blade angle axial flow fan, compressor or turbine as may be used in a turbojet or turbofan engine is described in this paper.
Abstract: A support, actuation, and balancing structure for rotor blades in a variable blade angle axial flow fan, compressor or turbine as may be used in a turbojet or turbofan engine. The supporting structure may include a shaft attached to each rotor blade and restrained from outward radial travel, under high centrifugal loading, by the inwardly facing surfaces of two spaced apart discs. Actuation may be provided by at least one fluid powered piston which controls blade angulation through rotation of one disc member relative to the other. Centrifugally controlled balancing means may also be provided to insure a uniform flow condition through the blades by maintaining the blades in a stable position of angulation upon disengagement of the actuator. Each rotor blade preferably includes a central axis about which the blade angle is varied, and which is in close proximity to the intersection of the leading edge of the blade with the blade tip.
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22 Jun 1971TL;DR: In this paper, a rotatable unison ring and linkage assembly are used to adjust the area of the throat defined between the next adjacent stator vanes, and a frame for circumferentially supporting the ring to prevent ovalizing of the ring and to provide accurate adjustment of the size of each of the throats areas.
Abstract: In an axial flow, fluid expansion engine having stator vanes adjustable, through a rotatable unison ring and linkage assembly, to vary the area of the throat defined between the next adjacent stator vanes, a frame for circumferentially supporting the unison ring to prevent ovalizing of the ring and, hence, provide accurate adjustment of the size of each of the throat areas.
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10 May 1971TL;DR: In this paper, an improved low pressure end diffuser for axial flow elastic fluid turbines, such as steam turbines, is disclosed, where a housing is provided on the outer fairing member of the diffuser, which housing defines a vented chamber.
Abstract: An improved low pressure end diffuser for axial flow elastic fluid turbines, such as steam turbines, is disclosed. A housing is provided on the outer fairing member of the diffuser, which housing defines a vented chamber. The outer fairing member is provided with openings communicating between the chamber and the annular area within the turbine casing adjacent the leading edges of the last stage turbine blades. Injection slots are formed in the outer fairing member to communicate between the chamber and the exhaust outlet downstream of the trailing edges of the last stage turbine blades. The injection slots are formed in such a manner that the fluid will flow from the chamber along a major portion of the inner surface of the outer fairing member whereby fluid boundary layer on the fairing surface will be accelerated to prevent separation of the boundary layer from the surface thereby resulting in improved diffuser performance.
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19 Apr 1971TL;DR: In this paper, a ring which is circumferentially expandable under centrifugal loading is proposed to constrain axial movement of individual blade members relative to the rotor axis.
Abstract: In an axial flow apparatus including stator and rotor elements, means are included for constraining axial movement of individual blade members relative to the rotor axis when the rotor is subjected to high centrifugal loading. The constraining means include a ring which is circumferentially expandable under centrifugal loading. Part of the outward radial force causing the expansion of the ring is directed axially against the blade members to provide an axially constraining force which is proportional to the centrifugal loading on the rotor. The invention herein described was made in the course of or under a contract or subcontract thereunder, (or grant) with the Department of the Army.
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TL;DR: In this article, a new boundary layer theory is developed which retains all elements of classical boundary layer theories, such as momentum thickness and wall shear stress, but introduces new concepts such as axial and tangential defect force thickness, a rotor exit-stator inlet "jump condition" and the importance of these concepts is demonstrated.
Abstract: The essential ingredient missing in existing prediction methods for the performance of multistage axial compressors is that which would account for the effect of end-wall boundary layers. It is, in fact, believed that end-wall boundary layers play a major role in compressor performance and the absence of an adequate theory represents a handicap to turbomachinery designers that might be likened to the handicap that designers of wings, for example, would face if Prandtl had not introduced the idea of a boundary layer. In this paper a new theory is developed which retains all elements of classical boundary layer theory; for example, we discuss variables such as momentum thickness and wall shear stress. However, the present theory introduces new concepts such as axial and tangential defect force thickness, a rotor exit-stator inlet “jump condition” and the importance of these concepts is demonstrated. Inherent in the derivation is an identification of the role of secondary flow and tip clearance flow. A proper means of matching the boundary layer calculations to conventional main stream calculations is suggested. Independent of empirical parametization it appears that the theory is capable of correctly modeling boundary layer blockage, losses, and end-wall stall. Near stall, the main stream-boundary layer interaction is very strong.
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01 Nov 1971
TL;DR: In this article, a throttling valve of the axial flow type for controlling fluid flow is described, which includes an expandable sleeve of resilient material having a toroidal configuration and positioned within a housing on a cage member.
Abstract: There is disclosed herein a throttling valve of the axial flow type for controlling fluid flow. The valve includes an expandable sleeve of resilient material having a toroidal configuration and positioned within a housing on a cage member. When the sleeve is in a rest position there is no fluid flow, but the sleeve can expand as a result of differential pressure acting thereon to allow flow in varying degrees between the sleeve and cage.
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01 Jan 1971
TL;DR: In this article, the root-mean-square (rms) displacement of a flexible rod or tube in axial flow is calculated in terms of beam natural frequency, damping factor, and intensity of the mean square spectral density of the pressure field in the low-frequency range.
Abstract: Many reactor and plant equipment components, such as fuel pins, control rods, and heat exchanger tubes, are long, slender, beam-like members which are exposed to nominally axial coolant flow. The flowing coolant represents a source of energy which can induce vibratory motion of these components. This design guide presents a relationship for calculating the root-mean-square (rms) displacement of a flexible rod or tube in axial flow. The relationship is based on the results of a parameter study and is valid for components that can be approximated as beams with either simply-supported or fixed-fixed ends. It is given in terms of beam natural frequency, damping factor, and intensity of the mean-square spectral density of the pressure field in the low-frequency range; all three are functions of mean axial flow velocity. Empirical expressions are developed for damping factor and intensity of the mean-square pressure spectrum. With these, an empirical equation for rms displacement is written which is in terms of known quantities and, therefore, provides a tool which can be used by designers. Since the equation is based on experiments involving a smooth rod in flow with minimal entrance effects, the predicted displacements should be interpreted and used with care. They are not conservative and, at best, will represent the minimum response to be expected. (auth)
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07 Dec 1971
TL;DR: A ground propulsion system for an aircraft includes a compressor driven by one of the main engines of the aircraft or an auxiliary engine, a turbine driven by the gaseous fluid supplied by the compressor and means including a gearbox coupling the turbine to a wheel of an aircraft as mentioned in this paper.
Abstract: A ground propulsion system for an aircraft includes a compressor driven by one of the main engines of the aircraft or an auxiliary engine, a turbine driven by the gaseous fluid supplied by the compressor and means including a gearbox coupling the turbine to a wheel of the aircraft.
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12 Oct 1971
TL;DR: In this article, a control means for preventing surging under changing or low flow conditions in a fluid compressor such as an air compressor which includes a rotor and diffuser and at least two stages of compression is disclosed.
Abstract: A control means for preventing surging under changing or low flow conditions in a fluid compressor such as an air compressor which includes a rotor and diffuser and at least two stages of compression is disclosed. Compressed air from the n+1 stage of the compressor is injected into the diffuser of the nth or lower stage of the compressor through a collection chamber which is disposed about the diffuser.
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TL;DR: In this article, a linearized analysis was performed to predict the velocity and static pressure redistribution in a distorted flowfield upstream of a low hub-tip ratio axial flow compressor as a function of the slope of the compressor pressure rise vs mass flow rate.
Abstract: A linearized analysis predicts the velocity and static pressure redistribution in a distorted flowfield upstream of a low hub-tip ratio axial flow compressor as a function of the slope of the compressor pressure rise vs mass flow rate characteristic. The attenuation of axial velocity 'distortion and the magnitude of the generated circumferential and radial velocities are found to increase with a steeper negative slope of the compressor characteristic. Analytical results indicate that the magnitude of the upstream flow redistribution is approximately halved when radial velocities are conserved within the compressor as compared to the results assuming that radial velocities are suppressed within the compressor. The test compressor suppressed internal radial velocities and data verified the appropriate analysis. The local slope of the distorted characteristic was found to be significantly less than the undistorted slope. Much of the difference between the local distorted slope and the undistorted compressor characteristic slope was attributable to rotor unsteady effects.
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01 Dec 1971TL;DR: A disc shaped bladed diaphragm structure for an axial flow turbine or compressor divided into two semicircular halves with radial grooves of a mating cylindrical cross-sectional shape and having one or more keying members disposed therein to provide a fluid seal thereacross is described in this paper.
Abstract: A disc shaped bladed diaphragm structure for an axial flow turbine or compressor divided into two semicircular halves with radial grooves of mating cylindrical cross-sectional shape and having one or more cylindrical keying members disposed therein to provide a fluid seal thereacross; and a method of making a diaphragm with the seal structure described above.
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TL;DR: In this paper, the sound radiated when inflow turbulence is present in axial flow fans has been investigated, and two noise radiating mechanisms can be identified: (i) interaction of turbulence with the rotor potential field results in a quadrupole-type volume source distribution, producing flow-interaction noise; (ii) impingement of turbulence on the blades results in dipole type (fluctuating force) surface source distribution.
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17 Jun 1971
TL;DR: In this paper, the inner blades are all inclined in the same direction and at the same angle in relation to the hub circumference, and the rim can expand freely under the effect of the centrifugal forces applied to it in operation by the outer blades.
Abstract: A gas turbine unit, in particular for an aircraft turbojet engine, comprises a fan rotating coaxially in front of an axial flow compressor. The fan has an outer blading ring fixed to a rim attached to a hub of the fan through an inner blading ring, and the inner blades are pivotally attached, at the root and tip ends respectively, to the hub and to the rim about pivot axes parallel to the axis of rotation of the hub. The inner blades are all inclined in the same direction and at the same angle in relation to the hub circumference. The rim can thus expand freely under the effect of the centrifugal forces applied to it in operation by the outer blading, without local distortion and without the application of any bending stresses to the inner blades.
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GMC1
TL;DR: In this paper, the tip portion of the impeller blades made of a porous material was used to allow some of the air to flow from the forward face of the blade through the blades to the rear face to promote more uniform velocity at the impeachmentller exit and thus reduce mixing losses and improve the slip factor of the compressor.
Abstract: A centrifugal compressor, which may otherwise be conventional, has the tip portion of the impeller blades made of a porous material to allow some of the air to flow from the forward face of the blade through the blades to the rear face to promote more uniform velocity at the impeller exit and thus reduce mixing losses and improve the slip factor of the compressor.
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15 Jun 1971TL;DR: A CIRCULAR SHROUDED ROTOR BLADE STRUCTURE for an ELASTIC FLUID AXIAL FLOW TURBINE or COMPRESSOR COMPRISING an ANNULAR ROW OF BLADES and AN ANNULAR SERIES of ARCUATE SEGMEN ASSOCIATED with the BLADes and ARRANGED in an end-to-end relationship.
Abstract: A CIRCULAR SHROUDED ROTOR BLADE STRUCTURE FOR AN ELASTIC FLUID AXIAL FLOW TURBINE OR COMPRESSOR COMPRISING AN ANNULAR ROW OF BLADES AND AN ANNULAR SERIES OF ARCUATE SEGMENST ASSOCIATED WITH THE BLADES AND ARRANGED IN END-TOEND RELATIONSHIP TO FORM A CONTINUOUS 360* SHROUD, WHEREIN EACH PAIR OF NEIGHBORING SHROUD SEGMENTS IS FIRMLY RIV- ETED TO A BLADE COMMON TO BOTH WITH EACH SEGMENT BEING RIVETED TO AT LEAST TWO BLADES. THERE MAY BE MORE THAN ONE ROW OF ARCUATE SEGMENTS FORMING THE SHROUD.
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17 Feb 1971
TL;DR: In this paper, a multistage axial flow compressor with stacked stages in an axial direction is described, each stage including a rotor and a stator, and the rotor and vane heights decrease from stage to stage in the direction of exiting fluid flow from the compressor.
Abstract: A multistage axial flow compressor with stacked stages in an axial direction, each stage including a rotor and a stator. The stages include vanes of identical vane geometry throughout the stator stages and the blades are of identical blade geometry throughout the several rotor stages. The blade and vane heights decrease from stage to stage in the direction of exiting fluid flow from the compressor. A constant hub diameter is maintained throughout the several rotor stages.
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01 Jan 1971
TL;DR: In this paper, the boundary layer behavior of axial-flow turbo-machine blading has been investigated and a new turbulent skin friction law for conditions of large positive pressure gradient is developed, and the problem of the minimum Reynolds number for turbulent flow under a pressure gradient was reexamined.
Abstract: Observations of the boundary layer behaviour on the blading of a single-stage axial-flow compressor are described : detailed measurements were carried out with hot wire, surface visualisation, and surface pitot tube techniques, and the presence of extensive regions of laminar flow was established. The behaviour of the laminar and turbulent boundary layers, and of separated laminar flow regions are reported. A new correlation is developed to describe the boundary layer transition behaviour, and the effects of pressure gradient, Reynolds number, and free stream turbulence on transition are discussed. A new turbulent skin friction law for conditions of large positive pressure gradient is developed, and the problem of the minimum Reynolds number for turbulent flow under a pressure gradient is re-examined. Various existing methods of calculating the turbulent boundary layer are examined and their success in predicting the boundary layer development on a stationary blade of the research compressor is evaluated. The application of the experimental results to the design and performance analysis of axial-flow turbo-machine blading is discussed. A family of surface velocity distributions giving unseparated flow over the suction surface of an axial-flow compressor blade is derived, and their computed performance is analysed. In conclusion, problems needing further investigation are outlined.