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A multistage vaneless counter-rotating turbine (MVCRT) eliminates vanes between rotors, which reduces the weight and size of the turbine and avoids viscous losses associated with vanes pronouncedly. An aircraft engine employing such a turbine would have greater thrust to weight ratio and smaller specific fuel consumption. This paper presents the aerodynamic design philosophy and performance analysis of the MVCRTs for gas turbine engines by a case study. The case is about a 1/2*4 turbine, which consists of a rotating frame and four rotors without any vanes between them. The first rotor and the third rotor are connected by a shaft to drive a compressor with a pressure ratio of 11.8, and the second rotor and the fourth rotor are connected by the rotating frame to deliver a total shaft power of around 2 MW. The stage loading of each rotor and flow axial acceleration of each duct are controlled to provide sufficient inlet swirls for their subsequent rotors. The stage work coefficients of each rotor are 0.95, 2.9, 1.4, and 1.0, respectively. Nonuniform radial circulation distributions are also used to maximize the turbine power output. Centrifugal forces in the outer rotor of the turbine are captured by carrying out a finite element analysis (FEA) to validate the aerodynamic design results. Three-dimensional viscous numerical results show that an adiabatic total-to-total efficiency of 91.47% with a pressure ratio of 9.8 at design condition is obtained and achieves the initial design objective very well. Entropy creation associated with the tip leakage and secondary flow is also illustrated for understanding the origins and effects of losses in the turbine. Pressure ratios and efficiency at the speed combinations of the 80% to 100% inner and outer rotor design speeds are discussed to reveal the turbine characteristics at off-design conditions.

Radial flow turbo machines have been used for a long time in a variety of applications such as turbochargers, cryogenics, auxiliary power units, and air conditioning of aircraft cabins. Hence numerous papers have been written on the design and performance of these machines. The only justification for yet another paper is that it would describe a unified approach for designing a single stage inward flow radial turbine comprising a rotor and the casing. The current turbine is designed to drive a direct-coupled permanent magnet high-speed alternator running at 60000 rpm and developing a maximum of 60 kW electrical power. The freedom of choice of the tip diameter and the tip width of the rotor that would be necessary for optimum isentropic efficiency of the turbine stage was restricted by the specified rotational speed and power output. Hence, an optimisation procedure was developed to determine the principal dimension of the rotor. The mean relative velocity in the rotor passages in the direction of the flow would be accelerated but flow velocity on the blade surfaces experiences a significant space rate of deceleration. The rate of deceleration can be controlled by means of a proper choice of the axial length of the rotor. A prescribed mean stream velocity distribution procedure was used to spread the rate of deceleration of the mean flow velocity along the meridional length of the flow passages. The nozzle-less volute casing was designed to satisfy the mass flow rate, energy and angular momentum equations simultaneously. This paper describes the work undertaken to design both the rotor and the casing. The work was motivated by the growing interest in developing gas turbine based hybrid power plant for road vehicles. The authors believe that the paper would lead to a stimulating discussion.Copyright © 2002 by ASME

This paper describes the time-varying aerodynamic interaction mechanisms that have been observed within a transonic high-pressure turbine stage; these are inferred from the time-resolved behavior of the rotor exit flow field. It contains results both from an experimental program in a turbine test facility and from numerical predictions. Experimental data was acquired using a fast-response aerodynamic probe capable of measuring Mach number, whirl angle, pitch angle, total pressure, and static pressure. A 3-D time-unsteady viscous Navier-Stokes solver was used for the numerical predictions. The unsteady rotor exit flowfield is formed from a combination of four flow phenomena: the rotor wake, the rotor trailing edge recompression shock, the tip-leakage flow, and the hub secondary flow. This paper describes the time-resolved behavior of each phenomenon and discusses the interaction mechanisms from which each originates. Two significant vane periodic changes (equivalent to a time-varying flow in the frame of reference of the rotor) in the rotor exit flowfield are identified. The first is a severe vane periodic fluctuation in flow conditions close to the hub wall and the second is a smaller vane periodic fluctuation occurring at equal strength over the entire blade span. These two regions of periodically varying flow are shown to be caused by two groups of interaction mechanisms. The first is thought to be caused by the interaction between the wake and secondary flow of the vane with the downstream rotor; and the second is thought to be caused by a combination of the interaction of the vane trailing edge recompression shock with the rotor, and the interaction between the vane and rotor potential fields.

This paper provides a physical interpretation of the mechanism of stagnation enthalpy and stagnation pressure changes in turbomachines due to unsteady flow, the agency for all work transfer between a turbomachine and an inviscid fluid. Examples are first given to illustrate the direct link between the time variation of static pressure seen by a given fluid particle and the rate of change of stagnation enthalpy for that particle. These include absolute stagnation temperature rises in turbine rotor tip leakage flow, wake transport through downstream blade rows, the influence on mixing losses of turbine wake behavior in downstream blade rows, and effects of wake phasing on compressor work input. Fluid dynamic situations are then constructed to explain the effect of unsteadiness, including a physical interpretation of how stagnation pressure variations are created by temporal variations in static pressure; in this it is shown that the unsteady static pressure plays the role of a time-dependent body force potential. It is further shown that when the unsteadiness is due to a spatial nonuniformity translating at constant speed, as in a turbomachine, the unsteady pressure variation can be viewed as a local power input per unit mass from this body force to the fluid particle at that point.Copyright © 2009 by ASME

The effect on rotor work of the phase of an upstream wake relative to the rotor is examined computationally and analytically for a transonic blade row. There can be an important impact on the time-mean performance when the time-dependent circulation of the shed vortices in the wake is phase-locked to the rotor position, as it occurs when there is strong interaction between the rotor static pressure field and the upstream vanes. The rotor work is found to depend on the path of the wake vortices as they travel through the blade passage; for the configurations examined, the calculated change in time-mean rotor work was approximately 3%. It is shown that the effect on work input can be analyzed in terms of the influence of the time-mean relative stagnation pressure nonuniformity associated with the unsteady (but phase-locked) wake vortex flow field, in that the changes in vortex path alter the location of the nonuniformity relative to the rotor. Lower pressure rise and work input occurs when the rotor blade is embedded in a region of low time-mean relative stagnation pressure than when immersed in a region of high relative stagnation pressure. In addition to the work changes, which are essentially two-dimensional effects, it is demonstrated that the location of the wake may affect the tip clearance flow, implying a potential impact on the pressure rise capability and rotor stability limits. Model calculations are presented to give estimates of the magnitude and nature of this phenomenon.

This paper describes the effect of tip clearance changes on the pressure at the case wall of a second-stage rotor. Wall shear distributions under the rotor tip are also presented. The results show low-pressure areas extending along the rotor suction side but lying away from the blade. Pressure contours indicate the tangential loading at the tip is lower than predicted by two-dimensional calculations; however, the predicted loading is observed between the lowest pressure’s path in the passage and the blade pressure side. The results suggest that a viscous or shearing layer, due to blade-to-wall relative motion, is generated on the blade side of the tip gap, which modifies the inviscid relative flow field and produces an unloading on the blade tip.

In this paper, three-dimensional multiblade row unsteady Navier-Stokes simulations at a hot streak temperature ratio of 2.0 have been performed to reveal the effects of rotor tip clearance on the inlet hot streak migration characteristics in high pressure stage of a Vaneless Counter-Rotating Turbine. The hot streak is circular in shape with a diameter equal to 25% of the high pressure turbine stator span. The hot streak center is located at 50% of the span and the leading edge of the high pressure turbine stator. The tip clearance size studied in this paper is 2.0mm (2.594% high pressure turbine rotor height). The numerical results indicate that the hot streak mixes with the high pressure turbine stator wake and convects towards the high pressure turbine rotor blade surface. Most of hotter fluid migrates to the pressure surface of the high pressure turbine rotor. Only a few of hotter fluid rounds the leading edge of the high pressure turbine rotor and migrates to the suction surface. The migration characteristics of the hot streak in the high pressure turbine rotor are dominated by the combined effects of secondary flow, buoyancy and leakage flow in the rotor tip clearance. The leakage flow trends to drive the hotter fluid towards the blade tip on the pressure surface and to the hub on the suction surface. Under the effect of the leakage flow, even partial hotter fluid near the pressure surface is also driven to the rotor suction surface through the tip clearance. Compared with the case without rotor tip clearance, the heat load of the high pressure turbine rotor is intensified due to the effects of the leakage flow. And the results indicate that the leakage flow effects trend to increase the low pressure turbine rotor inlet temperature at the tip region. The air flow with higher temperature at the tip region of the low pressure turbine rotor inlet will affect the flow and heat transfer characteristics in the downstream low pressure turbine.Copyright © 2008 by ASME

This paper presents an experimental investigation of a novel approach for controlling the rotor tip leakage and secondary flow by injecting cooling air from the stationary casing onto the rotor tip. It contains a detailed analysis of the unsteady flow interaction between the injected air and the flow in the rotor tip region and its impact on the rotor secondary flow structures. The experimental investigation has been conducted on a one-and-1/2-stage, unshrouded turbine, which has been especially designed and built for the current investigation. The turbine test case models a highly loaded, high pressure gas turbine stage. Measurements conducted with a two-sensor fast-response aerodynamic probe have provided data describing the time-resolved behavior of flow angles and pressures, as well as turbulence intensity in the exit plane of the rotor. Cooling air has been injected in the circumferential direction at a 30 deg angle from the casing tangent, opposing the rotor turning direction through a circumferential array of ten equidistant holes per rotor pitch. Different cooling air injection configurations have been tested. Injection parameters such as mass flow, axial position, and size of the holes have been varied to see the effect on the rotor tip secondary flows. The results of the current investigation show that with the injection, the size and the turbulence intensity of the rotor tip leakage vortex and the rotor tip passage vortex reduce. Both vortices move toward the tip suction side corner of the rotor passage. With an appropriate combination of injection mass flow rate and axial injection position, the isentropic efficiency of the stage was improved by 0.55 percentage points.

In radial turbines with variable nozzles, the flow downstream of the nozzles could be distorted by the leakage flow through the tip-clearance of the upstream nozzle vanes. To investigate the effects of flow distortion on the performance of turbine rotors, two rotors with different number of blades were tested for three types of distorted velocity distributions at the rotor inlet. In the case of the 20 blade rotor with moderate blade loading, the flow distortion at the rotor inlet had negligible effect on the rotor characteristics, and the measured data on the turbine performances agreed well with a prediction. Predictions were made with a conventional one-dimensional flow model applied to the rotor flow while a two-layer flow model was applied to the flow in the nozzle with clearance. In the case of the 10 blade rotor with heavy blade loading, however, the rotor performance was found to be sensitive to the inlet flow distortion and was considerably lower than the prediction.Copyright © 1989 by ASME

Between 1970 and 1974, ten variants of a supersonic axial compressor stage were designed and tested. These included two rotor configurations, three rotor tip clearances, addition of boundary-layer control consisting of vortex generators on both the outer casing and the rotor, and the introduction of slots in the stator vanes. Design performance objectives were a stage total pressure ratio of 3.0 with an isentropic efficiency of 0.82 at a tip speed of 1600 ft/s (488 m/s). The first configuration passed only 70 percent of design flow at design speed, achieving a stage pressure ratio of 2.25 at a peak stage isentropic efficiency of 0.61. The rotor was grossly separated. The tenth variant passed 91.4 percent of design flow at design speed, producing a stage pressure ratio of 3.03 with an isentropic efficiency of 0.75. The rotor achieved a pressure ratio of 3.59 at an efficiency of 0.87 under the same conditions. Major conclusions were that design tools available today would undoubtedly permit the original goals to be met or exceeded. However, the application for such a design is currently questionable because efficiency goals considered acceptable for most current programs have risen considerably from the level considered acceptable at the inception of this effort. Splitter vanes placed in the rotor permitted very high diffusion levels to be achieved without stalling. However, viscous effects causing three-dimensional flows violating the assumption of flow confined to concentric stream tubes were so strong that a geometry optimization does not appear practical without a three-dimensional, viscous analysis. Passive boundary-layer control in the form of vortex generators and slots does appear to offer some benefit under certain circumstances.