Semi-vaneless Space Research Articles

Numerical investigation of centrifugal compressor stability enhancement using vaned diffuser endwall contouring technology

To improve the stable operating range of centrifugal compressor stage, endwall contouring technology is introduced on the hub-side wall of the vaned diffuser for NASA CC3 centrifugal compressor as a promising technique to redistribute the flow field near the endwall in this paper. The main contents of this paper are listed as follows. First, a specific parameterized endwall contouring guideline is developed based on sinusoidal function and Bézier curve, and three factors are considered, namely, peak height, peak radial position, and frequency coefficient. Then, centrifugal compressor stages with baseline diffuser and contoured endwall diffusers are numerically investigated to reveal the influence of endwall contouring on the stable operating range. Results show that endwall contouring is an effective method to achieve stability enhancement with no evident reduction in stage performance on the design point. CEW13-12 has the most successful combination of endwall contouring parameters, and stall margin is increased by up to 40% for CEW13-12. Finally, the influence mechanism of endwall contouring on the stable operating range is discussed. At the near stall condition, the contouring endwall forces the fluid within the semi vaneless space to deflect toward the suction side to increase the momentum of low-energy fluid in the suction surface boundary layer and reduce the adverse pressure gradient on the suction side, which delay the development of flow separation on the vane leading edge near the shroud side to suppress the onset of rotating stall. At the near choke condition, the contouring endwall has more potential to increase the throat area, resulting in a decrease of throat blockage to improve choke margin. The results show that the endwall contouring technology is a reliable method to achieve stability enhancement of centrifugal compressor stage.

Numerical Study of the Improvement in Stability and Performance by Use of a Partial Vaned Diffuser for a Centrifugal Compressor Stage

The influence of different diffuser configurations on the flow stability and aerodynamic performance of a centrifugal compressor stage with a mass flow coefficient of 0.196 is numerically investigated. Research results show that the performance of a traditional full-height vaned diffuser (TVD) deteriorates rapidly, and a shroud-side partial vaned diffuser (SVD) displays better adaptability in off-design conditions. SVD can suppress the development of vortices generating at the diffuser leading-edge. Therefore, it can reduce the flow loss inside the stage and improve the flow stability of the stage at low mass flow rates. The unsteady analysis for TVD and SVD shows that the stall cell propagates at about 35.7% of impeller rotational speed in the semi-vaneless space and diffuser passages. Furthermore, the internal flow in TVD and SVD is studied by employing the proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD) methods. The flow loss and instability mechanism in the stage are consequently revealed more comprehensively.

Area Schedule Based Design of High-Pressure Recovery Radial Diffusion Systems

High-pressure ratio centrifugal compressors require advanced diffusion systems to achieve enhanced efficiencies set by future turbocharger applications. To address the shortcomings of the commonly used channel diffuser and airfoil cascade design perspectives, a streamtube based area schedule is adopted paying special attention to the diffuser entry region. It is shown that the diffusion in the semivaneless space, controlled chiefly by inlet flow angle and the vane suction side geometry, plays a key role in improving diffuser performance. Removing excess thickness from the suction side eliminates flow overspeed, increases effective diffusion length, and leads to higher pressure recovery at reduced stagnation pressure loss. The pressure side thickness distribution controls the channel area schedule. Thin leading edges (LEs) ensure smooth flow area transition into the channel and reduce the vane upstream influence, mitigating high-cycle fatigue related mechanical issues.

구심터빈의 노즐 내부 유동에 대한 시험 연구

Experimental study on the flow field inside the nozzle for radial turbine was performed. At design point, the pressure is high and the Mach number is low at the pressure side of the nozzle inlet semi-vaneless space as the flow turns through the nozzle vanes. As the flow accelerates through the nozzle passage to the throat the pressure level at the pressure and suction sides becomes similar. The flow continued accelerating from the throat to the inlet of turbine wheel and the pressure field became uniform in the circumferential direction in the vaneless space. In high expansion ratio condition, strong favorable pressure gradient band region occurred just after the throat in the semi-vaneless space in the circumferential direction and the pressure became uniform in the circumferential direction after this band. In low expansion ratio condition, core flow acceleration is dominant after the throat and this non-uniform pressure field reached to the inlet of turbine wheel.

Steady and Unsteady Flow Investigations on the Effect of Different Number of Diffuser Vanes

This paper presents a computational analysis of a high-speed centrifugal compressor stage where emphasis is focused on both the steady and unsteady flow investigations and on the effect of different number of diffuser vanes on the stage operating range. The simulations were carried out for a centrifugal stage with a backswept impeller and several diffusers with 11, 22 and 33 vanes. For the steady simulations, an averaging approach was used at the interface between the impeller and the diffuser. For the unsteady simulations, geometry scaling was used to deal with the problem of unequal pitch. As the number of diffuser vanes was increased, the pressure recovery coefficient in the semi-vaneless space at surge condition was found to reduce. Also, as the number of diffuser vanes was increased, the flow distortion became more pronounced in the vaneless and semi-vaneless space. It is proposed that this more distorted flow is harmful to the pressure recovery in the semi-vaneless space giving stages with more diffuser vanes narrower flow ranges. It was also found that the diffuser area ratio is not the dominant factor influencing the flow range when the number of diffuser vanes changes.

Experimental Investigation of Diffuser Hub Injection to Improve Centrifugal Compressor Stability

Results from a series of experiments to investigate whether centrifugal compressor stability could be improved by injecting air through the diffuser hub surface are reported. The research was conducted in a 4:1 pressure ratio centrifugal compressor configured with a vane-island diffuser. Injector nozzles were located just upstream of the leading edge of the diffuser vanes. Nozzle orientations were set to produce injected streams angled at −8, 0, and +8 degrees relative to the vane mean camber line. Several injection flow rates were tested using both an external air supply and recirculation from the diffuser exit. Compressor flow range did not improve at any injection flow rate that was tested, and generally diminished as injection rate increased. Compressor flow range did improve slightly at zero injection due to the flow resistance created by injector openings on the hub surface. Resistance and flow range both increased as the injector orientation was turned toward radial. Leading edge loading and semivaneless space diffusion showed trends that are similar to those reported earlier from shroud surface experiments that did improve compressor range. Opposite trends are seen for hub injection cases where compressor flow range decreased. The hub injection data further explain the range improvement provided by shroud-side injection and suggest that stability factors cited in the discussion of shroud surface techniques are valid. The results also suggest that a different application of hub-side techniques may produce a range improvement in centrifugal compressors.

Development of a performance prediction method for centrifugal compressor channel diffusers

A hybrid performance prediction method is proposed in the present study. A channel diffuser is divided into four subregions: vaneless space, semi-vaneless space, channel, and channel exit region. One-dimensional compressible core flow and boundary layer calculation of each region with an incidence loss model and empirical correlation of residuary pressure recovery coefficient of a channel predict the performance of diffusers. Three channel diffusers are designed and tested for validating the developed prediction method. The pressure distributions from an impeller exit to the channel diffuser exit are measured and discussed for various operating conditions from choke to nearly surge conditions. The strong non-uniform pressure distribution which is caused by impeller-diffuser interaction is obtained over the vaneless and semi-vaneless spaces. The predicted performance shows good agreement with the measured performance of diffusers at a design condition as well as at off-design conditions.

A Study on Pressure Distributions in a Centrifugal Compressor Channel Diffuser

Time averaged pressure distributions in a high-speed centrifugal compressor channel diffuser at design and off-design flow rates are investigated. Pressure distributions from the impeller exit to the channel diffuser exit are measured for various flow rates from choke to near surge condition, and the effects of operating condition are discussed. The strong non-uniformity in the pressure distribution is obtained over the vaneless space and semi-vaneless space caused by the impeller-diffuser interaction. As the flow rate increases, flow separation near the throat, due to large incidence angle at the vane leading edge, increases aerodynamic blockage and reduces the aerodynamic flow area downstream. Thus the minimum pressure location occurs downstream of the geometric throat, and it is named as the aerodynamic throat. And at choke condition, normal shock occurs downstream of this aerodynamic throat. The variation in the location of the aerodynamic throat is discussed.

A Discussion of the Factors Affecting Surge in Centrifugal Compressors

The process of surge in a centrifugal compressor has been studied and found to be dependent on a number of complex and often interrelated factors. A major factor defining surge in high-speed, vaned diffuser designs appears to be the flow in the semivaneless space. By developing work initiated by others it has been possible to propose some explanation of previously presented apparently conflicting data. In addition, a more complex mathematical model capable of assessing additional factors is proposed. Using this model good agreement with experimental surge has been obtained. This model takes into consideration pressure losses in the impeller, diffuser, and collector, and by using this model it was possible to identify the stalling elements that are responsible for overall compressor instability.