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 750ft∕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.

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.

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.

This thesis focuses on the use of aspiration on compressor blade design. The pressure ratio can be significantly increased by controlling the development of the blade and endwall boundary layers. This concept is 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 rotor tip speed of 750 ft/s. Aspiration equal to 0.5% of the inlet flow was applied to the blade surface of both the rotor and stator. Aspiration was also used on the endwall boundary layers. Detailed flowfield measurements are made behind the rotor and stator, and the ensemble-averaged data is compared with a 3-D, viscous analysis tool.The time-accurate flow measurements show a large blade to blade variation due to unsteady vortex shedding, which is not captured by conventional 3-D, viscous analysis tools. 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. 'Correction' of the experimental data using the model gives a better comparison with the 3-D, viscous analysis solution.In order to understand the possible benefits of aspiration, a meanline parameter study is performed over a range of rotor inlet Mach numbers, flow coefficients, and work coefficients. Viscous and shock losses are estimated for both conventional and aspirated stages. The results suggest that aspiration can have the largest impact on compressor performance at high stage pressure ratios.