Jack L. Kerrebrock

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

Heat Transfer in a Rotating Radial Channel With Swirling Internal Flow

This paper presents experimental results for heat transfer in swirling internal flow, obtained in two ways. A test rig simulated a rotating blade’s leading edge internal passage with heated walls and screw-shaped cooling swirl generated by flow introduced through discrete tangential slots. Spatially resolved variations of the surface heat transfer coefficients were measured in the rotating rig using an IR radiometer.A blade tested in the actual engine environment had similar geometry of the leading edge cooling passage. The blade surface temperatures were mapped in the engine with thermal paints and compared with a traditional convective cooling configuration. The data from the rotating rig and engine measurements are also compared with non-rotating heat transfer results obtained in the hot cascade using a traversing pyrometer at a realistic wall-to-coolant temperature ratio.The results are presented for realistic rotational numbers, ranging from 0 to 0.023, and for representative Reynolds number of 20,000 based on the channel diameter. The effect of Coriolis forces is evident with the change of direction of the rotation. A slight negative influence of the crossflow, which increased toward the outer radius of the channel, was recorded in the rig test results.The results presented will assist in better understanding of the screw-shaped swirl cooling technique, providing the next step toward the application of this highly-effective internal cooling method for the leading edges of turbine blades.© 1998 ASME

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

Propulsion System Requirements for Long Range, Supersonic Aircraft

This paper discusses the requirements for the propulsion system of supersonic cruise aircraft that are quiet enough to fly over land and operate from civil airports, have trans-pacific range in the order of 11,112km(6000nmi), and payload in the order of 4545kg(10,000lb.). It is concluded that the resulting requirements for both the fuel consumption and engine thrust/weight ratio for such aircraft will require high compressor exit and turbine inlet temperatures, together with bypass ratios that are significantly higher than typical supersonic-capable engines. Several technologies for improving both the fuel consumption and weight of the propulsion system are suggested. Some of these directly reduce engine weight while others, by improving individual component performance, will enable higher bypass ratios. The latter should therefore also indirectly reduce the bare engine weight. It is emphasized, however, that these specific technologies require considerable further development. While the use of higher bypass ratio is a significant departure from more usual engines designed for supersonic cruise, it is nonetheless considered to be a practical option for an aircraft of this kind.

Complementary Velocity and Heat Transfer Measurements in a Rotating Cooling Passage With Smooth Walls

An experimental investigation was conducted on the internal flowfield of a simulated smooth-wall turbine blade cooling passage. The square cross-sectioned passage was manufactured from quartz for optical accessibility. Velocity measurements were taken using Particle Image Velocimetry for both heated and non-heated cases. Thin film resistive heaters on all four exterior walls of the passage allowed heat to be added to the coolant flow without obstructing laser access. Under the same conditions, an infrared detector with associated optics collected wall temperature data for use in calculating local Nusselt number. The test section was operated with radial outward flow and at values of Reynolds number and Rotation number typical of a small turbine blade. The density ratio was 0.27. Velocity data for the non-heated case document the evolution of the coriolis-induced double vortex. The vortex has the effect of disproportionately increasing the leading side boundary layer thickness. Also, the streamwise component of the coriolis acceleration creates a considerably thinned side wall boundary layer. Additionally, these data reveal a highly unsteady, turbulent flowfield in the cooling passage. Velocity data for the heated case show a strongly distorted streamwise profile indicative of a buoyancy effect on the leading side. The coriolis vortex is the mechanism for the accumulation of stagnant flow on the leading side of the passage. Heat transfer data show a maximum factor of two difference in the Nusselt number from trailing side to leading side. A first-order estimate of this heat transfer disparity based on the measured boundary layer edge velocity yields approximately the same factor of two. A momentum integral model was developed for data interpretation which accounts for coriolis and buoyancy effects. Calculated streamwise profiles and secondary flows match the experimental data well. The model, the velocity data, and the heat transfer data combine to strongly suggest the presence of separated flow on the leading wall starting at about five hydraulic diameters from the channel inlet for the conditions studied.© 1998 ASME