Researchers are constantly pursuing technologies that will increase the performance of gas turbine engines. The aspirated compressor concept discussed here would allow the compression system to perform its task with about one-half of the compressor blades. To accomplish this, the researchers applied boundary layer control to the blades, casing, and hub. This method of boundary layer control consisted of removing small amounts of air from the main flow path at critical areas of the compressor. This bleed air could be used by other systems such as engine cooling or could be reinjected into lower pressure areas that require air for enhanced performance.
Left: Tip shrouded aspirated rotor with bleed slots on blade surfaces and bleed holes in shroud. Right: Aspirated rotor 50-percent wound with carbon fiber/epoxy matrix. Holes in the shroud are exit paths for bleeds taken from blades and shroud.
Long description of left figure.
Long description of right figure.
This effort was initiated by the Massachusetts Institute of Technology (MIT) in response to a solicitation from the Defense Advanced Research Projects Agency (DARPA) who sought to advance research in flow control technology. The NASA Glenn Research Center partnered with MIT (principal investigator), Honeywell Aircraft Engines (cycle analysis, structural analysis, and mechanical design), and Pratt & Whitney (cycle analysis and aeroanalysis) to conceptualize, design, analyze, build, and test the aspirated fan stage. The aerodesign and aeroanalysis of this fan stage were jointly executed by MIT and Glenn to minimize the amount of bleed flow needed and to maintain the highest efficiency possible (ref. 1). Mechanical design issues were complicated by the need to have a shrouded rotor with hollow blades, with rotor stress levels beyond the capabilities of titanium. The high stress issues were addressed by designing a shroud that was filament wound with a carbon fiber/epoxy matrix, resulting in an assembly that was strong enough to handle the high stresses. Both the rotor (preceding photographs) and stator (following photograph) were fabricated in two halves and then bolted together at the hub and tip, permitting the bleed passages to be machined into each half before assembly.
Aspirated stator with bleed slots on blade surfaces, bleed holes along the stator hub, and pitchwise bleed slot between blades.
Long description.
The aspirated fan stage was successfully tested at Glenn in December 2002 from 0 to 100 percent of the design speed and from 0 to 100 percent of the design mass flow rate. The stage pressure ratios and mass flow rates agreed very well with the pretest predicted values obtained from an in-house turbomachinery flow analysis code. This is a significant accomplishment given the complexity involved in the aerodynamic design of the blading and suction slot placements. The fan stage delivered a pressure ratio 50-percent higher than for conventional designs operating at the same speeds. This effort validates both the aerodesign and aeroanalysis codes used in this project. The aerodesign code known as MISES was developed by MIT, and the aeroanalysis code applied was APNASA, a Glenn in-house-developed code. Research in this area continues by NASA, academia, and industry.
Glenn contact: Edward Braunscheidel, 216-433-6298, Edward.P.Braunscheidel@nasa.gov
Author: Edward P. Braunscheidel
Headquarters program office: OAT
Programs/Projects: PR&T, UEET
Turbomachinery and Propulsion Systems
Last updated: January 21, 2005
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