This paper presents a new methodology for quantifying compressor endwall blockage and an approach, using this quantification, for defining the links between design parameters, flow conditions, and the growth of blockage due to tip clearance flow. Numerical simulations, measurements in a low-speed compressor, and measurements in a wind tunnel designed to simulate a compressor clearance flow are used to assess the approach. The analysis thus developed allows predictions of endwall blockage associated with variations in tip clearance, blade stagger angle, inlet boundary layer thickness, loading level, loading profile, solidity, and clearance jet total pressure. The estimates provided by this simplified method capture the trends in blockage with changes in design parameters to within 10 percent. More importantly, however, the method provides physical insight into, and thus guidance for control of, the flow features and phenomena responsible for compressor endwall blockage generation.

This paper presents an experimental and analytical study of the impact of forward swept rotors on tip-limited, low-speed, multi-stage axial compressors. Two different configurations were examined, one with strong tip-clearance flows and the other with more moderate levels. Evaluations were done at multiple rotor tip clearance levels to assess differences in clearance sensitivity. Both configurations are low-speed models of the rear stages of modern aircraft engine high pressure ratio compressors. Compared to conventionally stacked (radial) rotors, the forward swept blades demonstrated improvements in stall margin, efficiency and clearance sensitivity. The benefits were more pronounced for the configuration with stronger tip-clearance flows. Detailed flow measurements and 3-D viscous CFD analyses are used to investigate the responsible flow mechanisms. Forward sweep causes a spanwise redistribution of flow toward the blade tip and reduces the tip loading in terms of static pressure coefficient. This results in reduced tip-clearance flow blockage, a shallower (more axial) leakage/freestream interface angle and a smaller region of reversed flow in the clearance gap.

Model scale tests of modern 'open rotor' propulsor concepts that have potential for significant fuel burn reduction for aircraft applications were completed at NASA Glenn Research Center. The recent test campaign was a collaboration between NASA, FAA, and General Electric (GE). GE was the primary industrial partner, but other organisations were involved such as Boeing and Airbus who provided additional hardware for fuselage simulations. The open rotor is a modern version of the UnDucted Fan (UDF ® ) that was flight tested in the late 1980s through a partnership between NASA and GE. The UDF ® was memorable for its scimitar shaped propeller blades and its unique noise signature. Design methods of the time were not able to optimise for both high aerodynamic efficiency and low noise simultaneously. Contemporary CFD/CAA based design methods can produce open rotor blade designs that maintain efficiency with acceptable acoustic signatures. Tests of two generations of new open rotor designs were conducted in the 9' × 15' Low Speed Wind Tunnel and the 8' × 6' Supersonic Wind Tunnel starting in late 2009 and completed in early 2012. Aerodynamic performance and acoustic data were obtained for take-off, approach and cruise conditions in isolated and semi-installed configurations. Additional detailed flow diagnostic measurements and acoustic

This paper presents an experimental and analytical study of the impact of forward swept rotors on tip-limited, low-speed, multistage axial compressors. Two different configurations were examined, one with strong tip-clearance flows and the other with more moderate levels. Evaluations were done at multiple rotor tip clearances to assess differences in clearance sensitivity. Compared to conventionally stacked radial rotors, the forward swept blades demonstrated improvements in stall margin, efficiency and clearance sensitivity. The benefits were more pronounced for the configuration with stronger tip-clearance flows. Detailed flow measurements and three-dimensional viscous CFD analyses were used to investigate the responsible flow mechanisms. Forward sweep causes a spanwise redistribution of flow toward the blade tip and reduces the tip loading in terms of static pressure coefficient. This results in reduced tip-clearance flow blockage, a shallower (more axial) vortex trajectory and a smaller region of reversed flow in the clearance gap.

Building upon the successes of the UDF® program in the 1980’s, open rotor designs for high flight speed efficiency and low community noise have been developed at GE in collaboration with NASA and the FAA. Targeting a narrow body aircraft with 0.78 cruise Mach number, the cost-share program leveraged computational fluid dynamics (CFD), computational aero-acoustics (CAA), and rig scale testing to generate designs that achieved significant noise reductions well beyond what was attained in the 1980’s while substantially retaining cruise performance. This paper presents overall propeller net efficiency and acoustic assessments of GE’s modern open rotor designs based on measured rig data and the progression of the technology from the 1980’s through the present. Also discussed are the effects of aft rotor clipping, inter-rotor spacing, and disk loading. This paper shows how the two-phase design and scale model wind tunnel test program allowed for test results of the first design phase to feed back into the second design phase, resulting in 2–3% improvement in overall propeller net efficiency than the best efficiency design of the 1980’s while nominally achieving 15–17 EPNdB noise margin to Chapter 4 (when projected to full scale for a prescribed aircraft trajectory and installation). Accounting for trades and near term advancements, such a propulsion system is projected to meet the goal of 26% fuel burn reduction relative to CFM56-7B powered narrow body aircraft.

The relationship between turbomachinery blade circulation and tip clearance vortex circulation measured experimentally is examined using three-dimensional viscous flow computations. It is shown that the clearance vortex circulation one would measure is dependent on the placement of the fluid contour around which the circulation measurement is taken. Radial transport of vorticity results in the magnitude of the measured clearance vortex circulation generally being less than the blade circulation. For compressors, radial transport of vorticity shed from the blade tip in proximity to the endwall is the principal contributor to the discrepancy between the measured vortex circulation and blade circulation. Further, diffusion of vorticity shed at the blade tip toward the endwall makes it impossible in most practical cases to construct a fluid contour around the vortex that encloses all, and only, the vorticity shed from the blade tip. One should thus not expect agreement between measured tip clearance vortex circulation and circulation around the blade.

Rotor tip modeling fidelity, grid resolution, and near wall modeling have been examined to determine the requirements for an accurate prediction of the effects of large tip clearance in a low-speed axial compressor rotor. The effort, using a Reynolds-Averaged Navier-Stokes (RANS) solver, aimed to obtain the most accurate predictions from a three-dimensional, steady, single blade row simulation. A recently tested, modern low speed rotor, was used as the test geometry; the measured pressure rise characteristic as well as detailed data near stall was used to evaluate the ability of different modeling strategies to capture the correct flow structure. The leakage flow was quantified to show that a wide range of tip blockage could be obtained for different simulations of the same geometry and conditions. The results show that using a square tip and gridding to fully resolve the real tip gap was better able to capture the effects of loading on the leakage flow than either of the approximate models studied. Sufficient clustering near the casing to capture the shear layers was also found to be critical. While wall integration provided the best results in simultaneously improving the prediction of pressure rise characteristics and flow range, higher fidelity wall modeling and a casing y+ of approximately 3 were found to provide similar benefits.

This paper presents a new methodology for quantifying compressor endwall blockage and an approach, using this quantification, for defining the links between design parameters, flow conditions, and the growth of blockage due to tip clearance flow. Numerical simulations, measurements in a low speed compressor, and measurements in a wind tunnel designed to simulate a compressor clearance flow are used to assess the approach. The analysis thus developed allows predictions of endwall blockage associated with variations in tip clearance, blade stagger angle, inlet boundary layer thickness, loading level, loading profile, solidity and clearance jet total pressure. The estimates provided by this simplified method capture the trends in blockage with changes in design parameters to within 10%. More importantly, however, the method provides physical insight into, and thus guidance for control of, the flow features and phenomena responsible for compressor endwall blockage generation.