Annular Compressor Cascade Research Articles

Numerical investigation of a variable stator vane with nonuniform partial radial gaps in an annular compressor cascade

Abstract A variable stator vane (VSV) with nonuniform partial radial gaps is numerically investigated in an annular compressor cascade. Five adjusting angles (α = −5.5°, −5°, 0°, 5° and 10°) are chosen to simulate different working conditions. The VSV is adjusted in negative direction means that the stagger angle increases and the VSV is more closed. Since the VSV is installed in an annular cascade, the heights of the partial gaps are nonuniform and reaches 3.8% of the span. Results show that as the VSV is adjusted negatively, the total pressure loss rises by 20.3% and a huge area of corner separation appears. The outflow angle is also more distorted along radial direction. Comparisons with a fixed-configuration stator show that although the VSV would cause higher loss, it indeed plays an important role in adjusting the outflow angle.

Spanwise Variations in Aerodynamic Damping of an Oscillating Annular Compressor Cascade

Aeroelastic instability in turbomachinery is one of the severe problems limiting its performance. Of particular interest is the phenomenon of stall flutter in compressors, which occurs when blades encounter large incidence flows. With the objective to understand the aerodynamics associated with stall flutter, the present experimental study is conducted on a subsonic annular compressor cascade. Five consecutive blades are oscillated, with the central blade instrumented for unsteady surface pressure measurements. The flow incidence is established using inlet guide vanes to provide three inflow profiles, representative of nominal, near-stall, and poststall incidences. With the blades oscillating at specific interblade phase angles and reduced frequencies, the stability of the cascade is determined globally and locally in terms of aerodynamic damping. The near-hub flow is distinctive due to the thickened boundary layer, resulting in an irregular variation of pressure with low magnitudes, eventually approaching negative stability. At larger incidences, due to the low-amplitude fluctuations in pressure at all spans, the overall stability is reduced. The present work also describes and rationalizes the nature of the dependence of blade unsteady pressure distribution on reduced frequency and incidence.

Experimental Study of Aerodynamic Damping of an Annular Compressor Cascade With Large Mean Incidences

This study addresses flutter that can occur in compressors when operating at high relative incidence. Studies are performed on a subsonic annular compressor cascade containing a sector of blades that can be subjected to controlled torsional oscillation. Measurements taken on the centrally located blade are used to study the unsteady surface pressures developed. Three large mean incidences are considered to characterize the aeroelastic performance. Aerodynamic damping is calculated for each test condition and its variation due to interblade phase angle (IBPA), reduced frequency, and incidence is studied. The source of stability or instability is traced to the nature of unsteady pressures. When the incidence is below the static-stall limit, an increasing incidence is found to exhibit aeroelastically more stable performance, whereas beyond the limit, the stability worsens. For the latter, the amount of improvement in stability by increasing reduced frequency is less compared to the former and its variation with IBPA is not as regular owing to the associated large uncertainty. The nonlinearity effects were found to be relatively higher for this case, especially from the aft region of the suction surface. It is also found that the phase of the local fluctuating pressure and its location on the chord has a decisive influence on the aerodynamic damping and its trends with respect to various parameters are discussed. The results are expected to be useful in the assessing aerodynamic damping trends in relation to the pressure phase variations in specific regions along the chord.

Passive flow control of the corner separation in an annular compressor cascade with a micro-blade

Corner separation contributes greatly to the loss and the passage blockage in a compressor stage. In order to mitigate the corner separation and improve the aerodynamic performance of compressors, a novel passive flow control method, an off-surface micro-blade installed upstream of the separation onset location, was proposed. A numerical investigation has was performed in an annular compressor cascade to assess the control effectiveness of the micro-blade. The results show that the location of the micro-blade affects the control effect significantly. The application of the well-designed micro-blade enhances the diffusion capacity considerably under the inflow incidence from −2° to +10°, accompanied by a slight loss reduction at some particular incidences. Detailed analysis of the predicted flow field was carried out to understand the underlying mechanism. It indicates that a “jet” forms upstream of the separation onset location with the application of the micro-blade. The formation of the jet reduces the thickness of boundary layer on the suction surface and builds a “jet barrier” near the endwall to hinder the accumulation of the low momentum fluid. The influence of the incidence was also investigated. It is concluded that the incidence increase has both positive and negative influences on the control effect of the micro-blade. As a result, the performance of the micro-blade is sensitive to the variation of inlet incidence.

Investigation of Leading Edge Tubercles with Different Wavelengths in an Annular Compressor Cascade

Abstract Humpback whales possess bumpy tubercles on the leading edge of their flippers. Due to these leading edge tubercles, whales are able to produce high degree of maneuverability. Inspired by the flippers, this paper applies sinusoidal-like tubercles to the leading edge of blades in an annular compressor cascade, and presents a numerical investigation to explore the effects of tubercles with the aim of controlling the corner separation and reducing losses. Steady 3D RANS simulations are performed to investigate the aerodynamic performance and behavior of the corner separation in compressor cascades with and without leading edge tubercles. A crucial geometry parameter of the tubercles, wavelength, is varied to obtain different configurations. Results show that a smaller wavelength (more wave number) corresponds to a larger loss reduction and the maximum loss reduction reaches to 46.0%. Also, it is found that leading edge tubercles result in a stall delay and the maximum stall angle improvement reaches to 28.1%. Flow visualizations show that leading edge tubercles could induce the formation of counter-rotating streamwise vortices. The interaction between the streamwise vortices and corner separation is thought to be the primary flow mechanism generated by leading edge tubercles in an annular compressor cascade.

Flow Control of Annular Compressor Cascade by Synthetic Jets

An experimental investigation conducted in a stationary annular cascade wind tunnel demonstrated that unsteady flow control using synthetic jets (zero mass flux) could effectively reduce flow separation in the axial compressor cascade. The synthetic jets driven by speaker were introduced through the casing radially into the flow-field just adjacent to the leading edge of the compressor cascade. The experimental results revealed that the aerodynamic performance of the compressor cascade could be improved amazingly by synthetic jets and the maximum relative reduction of loss coefficient was up to 27.5%. The optimal analysis of the excitation frequency, excitation location was investigated at different incidences. In order to obtain detailed information on flow-field structure, the digital particle image velocimetry (DPIV) technique was adopted. The experimental results indicated that the intensity of wake vortices became much weaker and streamlines became smoother and more uniform with synthetic jets.

Computation Of The NTUA/LTT CompressorAnnular-cascade Flow Using TheMulti-dimensional Upwind Residual DistributionMethod

This paper extends to a rotating co-ordinate system the three-dimensionalsymmetric form, in space and time, of the Euler equations, previously reported by the authors, in an inertial frame of reference. This form of the Euler equations is discretized in space using the multi-dimensionalupwindresidualdistributionmethod. The integration of the cell residual and the evaluation of the distribution matrices is performed using the Roe’s parameter vector and a mean weighted square radius. An implicit time step is adopted to reduce the computational time required to achieve the steady-state solution. Numerical results are presented for the flow through the NTUA/LTT annular-compressor cascade in a rotating frame of reference.

A Correlation-Based Transition Model Using Local Variables—Part II: Test Cases and Industrial Applications

A new correlation-based transition model has been developed, which is built strictly on local variables. As a result, the transition model is compatible with modern computational fluid dynamics (CFD) methods using unstructured grids and massive parallel execution. The model is based on two transport equations, one for the intermittency and one for the transition onset criteria in terms of momentum thickness Reynolds number. The proposed transport equations do not attempt to model the physics of the transition process (unlike, e.g., turbulence models), but form a framework for the implementation of correlation-based models into general-purpose CFD methods. Part I of this paper (Menter, F. R., Langtry, R. B., Likki, S. R., Suzen, Y. B., Huang, P. G., and Völker, S., 2006, ASME J. Turbomach., 128(3), pp. 413–422) gives a detailed description of the mathematical formulation of the model and some of the basic test cases used for model validation. Part II (this part) details a significant number of test cases that have been used to validate the transition model for turbomachinery and aerodynamic applications, including the drag crisis of a cylinder, separation-induced transition on a circular leading edge, and natural transition on a wind turbine airfoil. Turbomachinery test cases include a highly loaded compressor cascade, a low-pressure turbine blade, a transonic turbine guide vane, a 3D annular compressor cascade, and unsteady transition due to wake impingement. In addition, predictions are shown for an actual industrial application, namely, a GE low-pressure turbine vane. In all cases, good agreement with the experiments could be achieved and the authors believe that the current model is a significant step forward in engineering transition modeling.

Hub Wall Rotation Influence on High-Speed Compressor Cascade Performance

The ine uence of the magnitude of the relative wall speed on the performance of an annular compressor cascade with hub clearance is examined. Five-hole probe measurements, conducted at the inlet and outlet of the cascade, are used to derive blade performance characteristics, in the form of loss and turning distributions. Characteristics are presented in the form of circumferentially mass-averaged proe les, whereas distributions on the exit plane provide information useful to interpret the performance of the blading. Four different rotational speeds of the hub have been examined, giving the possibility to observe the dependence of performance characteristics on hub rotational speed. Increasing the rotational speed is found to improvethe performance of the cascade by decreasing losses in the clearance region, while it affects the e ow in the entire passage.

On the Navier–Stokes Calculation of Separation Bubbles With a New Transition Model

Laminar separation bubbles are commonly observed on turbomachinery blades and therefore require effective methods for their prediction. Therefore, a newly developed transition model by Gostelow et al. (1996) is incorporated into an upwind-biased Navier–Stokes code to simulate laminar—turbulent transition in the boundary layer. A study of the influence of the two adjustable parameters of the model, the transition onset location and the spot generation rate, is conducted and it is found that it can predict laminar separation bubbles, measured on a NACA 0012 airfoil. Additional results are presented for separation bubbles in an annular compressor cascade.

Experimental investigation of the three-dimensional flow in an annular compressor cascade at large incidence

This paper presents a detailed experimental investigation concerning the influence of blade loading (incidence) on the three-dimensional flow in an annular compressor cascade. The data are acquired at four incidence angles under low Mach number and low Reynolds number conditions. Experimental techniques include the oil-film visualization on the profile and the endwall surfaces, the laser-sheet visualization of the flow field inside the blade passage, and the measurement by radial-circumferential traverses using a seven-hole probe. The behavior and nature of the three-dimensional flow with severe separations inside the blade passage and at the exit are obtained. The distributions of the total pressure loss, static pressure, velocity and outflow angle are also given. These results are valuable for establishing the physical model of the three-dimensional complex flow in axial compressors and for examining the computational procedures.

Experimental and Numerical Investigation of Three-Dimensional Viscous Flows and Vortex Motion Inside an Annular Compressor Blade Row

A detailed experimental and numerical investigation was carried out to examine the three-dimensional flow field, secondary flows, and vortex motion in an annular compressor cascade. Various flow visualizations near the blade surface and endwalls, wall static pressure and loss measurements, as well as hot-film and hot-wire measurements inside the blade boundary layers were performed at various flow rates to understand the complex flow phenomena. A Reynolds-averaged Navier–Stokes equation was solved to investigate the flow numerically. The detailed comparison between measurement and numerical prediction indicates that the complex three-dimensional flow phenomena (corner stall, vortex motion, radial mixing, etc.) are very well predicted with the numerical method.

Three-Dimensional Wake Decay Inside of a Compressor Cascade and Its Influence on the Downstream Unsteady Flow Field: Part I—Wake Decay Characteristics in the Flow Passage

A measuring technique based on multisensor hot-wire anemometry has been developed to determine the unsteady three-dimensional velocity vector and the structure of turbulent flows. It has then been applied to the passage and the exit flow of an annular compressor cascade, which is periodically disturbed by the wakes of a cylinder rotor, located about 50 percent of blade chord upstream. In part I of this paper the decay of the rotor wakes will be described first without stator and secondly through a stator passage. The time-dependent turbulent flow field downstream of this stator is discussed in Part II. The rotor wakes have a major influence on the development of three-dimensional separated regions inside the compressor cascade, and this interaction will be addressed in both parts of this paper.

Three-Dimensional Wake Decay Inside of a Compressor Cascade and Its Influence on the Downstream Unsteady Flow Field: Part II—Unsteady Flow Field Downstream of the Stator

A measuring technique based on multisensor hot-wire anemometry has been developed to determine the unsteady three-dimensional velocity vector and the structure of turbulent flows. It then has been applied to the passage and the exit flow of an annular compressor cascade, which is periodically disturbed by the wakes of a cylinder rotor, located about 50 percent of blade chord upstream. In Part I of this paper the decay of the rotor wakes has been described first without stator and secondly through a stator passage. The time-dependent turbulent flow field downstream of this stator is discussed in Part II of this paper. The rotor wakes have a major influence on the development of three-dimensional separated regions inside the compressor cascade, and this interaction will be addressed in both parts of the paper.

Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 1—Quasi-Steady Flow Field and Comparison With Steady-State Data

An experimental study of three-dimensional flow field in an annular compressor cascade with an upstream rotor has been carried out at four different incidences to the stator blade. Blade boundary layers and the three-dimensional flow field at the exit are surveyed using a hot-wire sensor and a five-hole probe, respectively. The data on the blade boundary layer, passage flow, and separated corner flow are presented. The upstream rotor wake has a major influence on the transition, laminar separation bubble, extent of wall/corner flow separation, aerodynamic losses, secondary flow, and three-dimensional flow inside the passage. A detailed interpretation of the effects of upstream wakes on the entire passage flow is presented and compared with the data in the absence of a rotor.

Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 2—Unsteady Flow and Pressure Field

An experimental study of the unsteady three-dimensional flow and pressure field in an annular compressor cascade with an upstream rotor has been carried out at several incidences to the stator blade. The distributions of the unsteady pressures at the blade surfaces are measured using fast response Kulite sensors. The unsteady blade boundary layers and the passage flow are measured with a hot-wire sensor. A detailed interpretation of the magnitude of unsteady pressures, phase angle differences, unsteady blade boundary layers, and wake transport through the stator passage is presented and analyzed in the paper. The unsteady pressures are found to be dominant near the blade leading edge. Substantially higher pressure fluctuations occur in this region as well as on the edge of the corner flow separation region.

Closure to “Discussions of ‘Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 2—Unsteady Flow and Pressure Field’” (1990, ASME J. Turbomach., 112, pp. 688–689)

Closure to “Discussions of ‘Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 2—Unsteady Flow and Pressure Field’” (1990, ASME J. Turbomach., 112, pp. 688–689)

Discussion: “Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 2—Unsteady Flow and Pressure Field” (Schulz, H. D., Gallus, H. E., and Lakshminarayana, B., 1990, ASME J. Turbomach., 112, pp. 679–688)

Discussion: “Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 2—Unsteady Flow and Pressure Field” (Schulz, H. D., Gallus, H. E., and Lakshminarayana, B., 1990, ASME J. Turbomach., 112, pp. 679–688)

Discussion: “Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 2—Unsteady Flow and Pressure Field” (Schulz, H. D., Gallus, H. E., and Lakshminarayana, B., 1990, ASME J. Turbomach., 112, pp. 679–688)

Discussion: “Three-Dimensional Separated Flow Field in the Endwall Region of an Annular Compressor Cascade in the Presence of Rotor-Stator Interaction: Part 2—Unsteady Flow and Pressure Field” (Schulz, H. D., Gallus, H. E., and Lakshminarayana, B., 1990, ASME J. Turbomach., 112, pp. 679–688)

Experimental Investigation of the Three-Dimensional Flow in an Annular Compressor Cascade

A detailed experimental investigation was carried out to examine the influence of blade loading on the three-dimensional flow in an annular compressor cascade. Data were acquired over a range of incidence angles. Included are airfoil and endwall flow visualization, measurement of the static pressure distribution on the flow passage surfaces, and radial-circumferential traverse measurements. The data indicate the formation of a strong vortex near the rear of the blade passage. This vortex transports low-momentum fluid close to the hub toward the blade suction side and seems to be partly responsible for the occurrence of a hub corner stall. The effect of increased loading on the growth of the hub corner stall and its impact on the passage blockage are discussed. Detailed mapping of the blade boundary layer was done to determine the loci of boundary layer transition and flow separation. The data have been compared with results from an integral boundary layer method.