Ali Merchant

Experimental investigation of a high pressure ratio aspirated fan stage

The experimental investigation of an aspirated fan stage designed to achieve a pressure ratio of 3.4:1 at 1500 ft/s is presented in this paper. The low-energy viscous flow is aspirated from diffusion-limiting locations on the blades and flowpath surfaces of the stage, enabling a very high pressure ratio to be achieved in a single stage. The fan stage performance was mapped at various operating speeds from choke to stall in a compressor facility at fully simulated engine conditions. The experimentally determined stage performance, in terms of pressure ratio and corresponding inlet mass flow rate, was found to be in good agreement with the 3D viscous computational prediction, and in turn close to the design intent. Stage pressure ratios exceeding 3:1 were achieved at design speed, with an aspiration flow fraction of 3.5% of the stage inlet mass flow. The experimental performance of the stage at various operating conditions, including detailed flowfield measurements, are presented and discussed in the context of the computational analyses. The stage performance and operability at reduced aspiration flow rates at design and off-design conditions are also discussed.

A FAMILY OF DESIGNS FOR ASPIRATED COMPRESSORS

The performance of compressors can be enhanced by the judicious removal of the viscous boundary layer fluid from the flow path. Removal of the boundary layer fluid just prior to or in a region of rapid pressure rise, either at shock impingement or more generally at the point of rapid pressure rise on the suction surface of the blade, can give significant increases in the diffusion and therefore increase the work done per stage for a given blade speed. It also provides a thermodynamic benefit by removing the high-entropy fluid from the flow path. Design studies have been done using quasi 3-D viscous and 3-D Euler computational tools on a family of fan stages of varying tip speed that lake advantage of such viscous fluid removal. One stage in this family is a low tip speed fan stage designed to produce a pressure ratio of 1.5 at a tip speed of 700 ft/sec. Fan noise reductions resulting from the decrease in tangential Mach number, without sacrificing total pressure ratio, could make this design attractive for the fan of medium-bypass ratio engines. Another stage in the family would produce a total pressure ratio of 2.0 at a tip speed of 1000 ft/sec and could be very attractive as a fan stage on a lower bypass ratio engine or as a first stage of a low speed core compressor. The final stage in the family would achieve a pressure ratio of more than 3.0 at a tip speed of 1500 ft/sec and could be very attractive as a first stage of a core compressor, or as a fan for a military engine. A design for the suction passages to deal with the fluid removal has been completed for an experimental version of the 1.5 pressure ratio design. A tip shroud allows bleeding of the tip surface boundary layer from the rotor, and carries the fluid removed from the blade surfaces through the tip. One of these stages will be tested in the MIT Blowdown Compressor, serving a dual purpose: as a validation of the computational design process and as a test of the concept of aspirated compressors.Copyright

Experimental Investigation of a Transonic Aspirated Compressor

The experimental investigation of a transonic aspirated stage demonstrating the application of boundary layer aspiration to increase stage work is presented. The stage was designed to produce a pressure ratio of 1.6 at a tip speed of 750 ft/s resulting in a stage work coefficient of 0.88. The primary aspiration requirement for the stage is a bleed fraction 0.5% of the inlet mass flow on the rotor and stator suction surfaces. Additional aspiration totaling 2.8% was also used at shock impingement locations and other locations on the hub and casing walls. Detailed rotor and stator flow field measurements, which include time-accurate and ensemble-averaged data, are presented and compared to three-dimensional viscous computational analyses of the stage. The stage achieved a peak pressure ratio of 1.58 and through-flow efficiency of 90% at the design point. In addition, the stage demonstrated good performance with an aspiration lower than the design requirement, and a significant off-design flow range below that predicted by the computational analysis.

Design and Test of an Aspirated Counter-Rotating Fan

The design and test of a two-stage, vaneless, aspirated counter-rotating fan is presented in this paper. The fan nominal design objectives were a pressure ratio of 3:1 and adiabatic efficiency of 87%. A pressure ratio of 2.9 at 89% efficiency was measured at the design speed. The configuration consists of a counter-swirl-producing inlet guide vane, followed by a high tip speed (1450 ft/s) nonaspirated rotor and a counter-rotating low speed (1150 ft/s) aspirated rotor. The lower tip speed and lower solidity of the second rotor result in a blade loading above conventional limits, but enable a balance between the shock loss and viscous boundary layer loss; the latter of which can be controlled by aspiration. The aspiration slot on the second rotor suction surface extends from the hub up to 80% span. The bleed flow is discharged inward through the blade hub. This fan was tested in a short duration blowdown facility. Particular attention was given to the design of the instrumentation to measure efficiency to 0.5% accuracy. High response static pressure measurements were taken between the rotors and downstream of the fan to determine the stall behavior. Pressure ratio, mass flow, and efficiency on speed lines from 90% to 102% of the design speed are presented and discussed along with comparison to computational fluid dynamics predictions and design intent. The results presented here complement those presented earlier for two aspirated fan stages with tip shrouds, extending the validated design space for aspirated compressors to include designs with conventional unshrouded rotors and with inward removal of the aspirated flow.

Aerodynamic Design and Analysis of a High Pressure Ratio Aspirated Compressor Stage

The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of a unique aspirated compressor stage which achieves a total pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated stage was designed using an axisymmetric through-flow code coupled with a quasi three-dimensional cascade plane code with inverse design capability. Validation of the completed design was carried out with three-dimensional Navier-Stokes calculations. Spanwise slots were used on the rotor and stator suction surfaces to bleed the boundary layer with a total suction requirement of 4% of the inlet mass flow. Additional bleed of 3% was also required on the hub and shroud near shock impingement locations. A three-dimensional viscous evaluation of the design showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The three-dimensional viscous analysis predicted a mass averaged total pressure ratio of 3.7 at an isentropic efficiency of 93% for the rotor, and a mass averaged total pressure ratio of 3.4 at an isentropic efficiency of 86% for the stage.Copyright

Aerodynamic Design and Performance of Aspirated Airfoils

The impact of boundary layer aspiration, or suction, on the aerodynamic design and performance of turbomachinery airfoils is discussed in this paper. Aspiration is studied first in the context of a controlled diffusion cascade, where the effect of discrete aspiration on loading levels and profile loss is computationally investigated. Blade design features which are essential in achieving high loading and minimizing the aspiration requirement are described. Design studies of two aspirated compressor stages and an aspirated turbine exit guide vane using three dimensional Navier-Stokes calculations are presented. The calculations show that high loading can be achieved over most of the blade span with a relatively small amount of aspiration. Three dimensional effects close to the endwalls are shown to degrade the performance to varying degrees depending on the loading level.

Design, Analysis, Fabrication and Test of an Aspirated Fan Stage

A fan stage designed by means of a MISES-based quasi-3D approach (Youngren and Drela, 1991), for a pressure ratio of 1.6 at a tip Mach number of 0.7, has been analyzed by viscous 3D CFD, fabricated and tested in the MIT Blowdown Compressor. The design incorporates a rotor tip shroud and boundary layer removal on the suction surfaces of the rotor and stator and at other critical locations. The fully viscous 3D analysis enabled final detailing of the design. In tests, the stage has met its design objectives, producing the design pressure ratio of 1.6 at design speed. The mass flow removed totaled 4.7%, approximately 1.0% through slots on the suction surface of the rotor and stator, and the remainder distributed over the rotor shroud and stator hub. The measured adiabatic efficiency of the rotor for the throughflow was 96% at the design point and that for the stage was 90%. This paper presents the design, the results of the analysis and the experimental stage performance both at design and at some off-design conditions.Copyright

Experimental Investigation of an Aspirated Fan Stage

This paper addresses the use of viscous flow control by suction to improve compressor stage performance. The pressure ratio can be significantly increased by controlling the development of the airfoil and endwall boundary layers. This concept has been validated through an aspirated fan stage experiment performed in the MIT Blowdown Compressor Facility. The fan stage was designed to produce a pressure ratio of 1.6 at a throughflow adiabatic efficiency of 89% at a corrected rotor tip speed of 750 ft/s (229 m/s). Aspiration was applied to the airfoil surface of both the rotor and stator at a design suction rate of 0.5% of the inlet flow. Aspiration was also used on the endwall boundary layers. The measured performance of the stage agrees well with the design intent and predicted performance. An incompressible, vortex shedding model calibrated to the experimental data shows that the vortex shedding induces radial flows that redistribute flow properties in the spanwise direction.Copyright

A CAD-Based Blade Geometry Model for Turbomachinery Aero Design Systems

A CAD-centric approach for constructing and managing the blade geometry in turbomachinery aero design systems is presented in this paper. Central to the approach are a flexible CAD-based parametric blade model definition and a set of CAD-neutral interfaces which enable construction and manipulation of the blade solid model directly inside the CAD system’s geometry kernel. A bottleneck of transferring geometry data passively via a file-based method is thus eliminated, and a seamless integration between the CAD system, aero design system, and the larger design environment can be achieved. A single consistent CAD-based blade model is available at all stages of the aero design process, forming the basis for coupling the aero design system to the larger multi-disciplinary design environment. The blade model construction is fully parameterized so that geometry updates can be accurately controlled via parameter changes, and geometric sensitivities of the model can be easily calculated for multidisciplinary interaction and design optimization. A clear separation of the parameters that control the three-dimensional shape of the blade (such as lean and sweep) from the parameters that control the elemental profile shape allows any blade profile family or shape definition to be utilized. The blade model definition, construction interface, and implementation approach are described. Applications illustrating solid model construction, parametric modification and sensitivity calculation, which are key requirements for automated aerodynamic shape design, are presented.Copyright

Optimization of a 3-Stage Booster: Part 2—The Parametric 3D Blade Geometry Modeling Tool

A parametric approach for blade geometry design has been developed to obtain 3D blade models. The geometry of the blade is defined by a basic set of parameters that are first obtained from an axisymmetric solver. These parameters include the leading edge meridional coordinates, flow angles, axial chord, and the meridional coordinates of streamlines. Other parameters such as thickness to chord ratio need to be defined. Using these parameters the 2D airfoils are created and are stacked radially using one of the many multiple options that define the stacking axes from several additional parameters. The tool produces the desired number of 2D sections in a normalized coordinate system. Each blade section is then transformed to a 3D Cartesian coordinate system. Using Unigraphics-NX (CAD package), these sections are lofted and a 3D blade model is obtained. Parametric update of the spline points defining the 3D blade sections results in new blade shapes without going directly back into the CAD system. The importing of the geometry into a CFD solver, and a finite element solver to determine mode shapes and stresses is demonstrated. Full details of the blade procedure is presented for a 3-Stage Booster design. This parametric approach for defining blade geometry and how it lays a groundwork for a high-fidelity optimization procedure is described.Copyright