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Experimental evaluation of outer case blowing or bleeding of single stage axial flow compressor, and performance tests using distorted or undistorted inlet flow
Boundary layer bleed techniques for high lift stator blades, with hub diffusion factor stator performance data
Outer case bleeding or blowing effects on casing insert of single stage axial flow compressor with distorted inlet flow
Boundary layer bleed techniques for high lift compressor stator blades
Performance tests on slotted hub diffusion factor stator with boundary layer bleed
Evaluation of outer case blowing or bleeding of single stage axial flow compressor
Numerical study on controlling dynamic properties of supersonic inlet using bypass bleed
Determining feasibility of increasing stator blade loading of single stage compressor rig by bleeding or blowing suction surface boundary layer
Velocity profile control tests of diffuser wall bleed to control combustor inlet airflow distribution
Turbulent incompressible air flow in-stream static pressure fluctuations measurement, describing bleed type pressure transducer theory, design and operational characteristics
Increased guide-vane turning resulted in poorer overall performance, the decrease being greatest at the highest rotor speed. Rotating stall originating at the tips of the first stage correlated with the knee in the stall-limit line. Increasing guide-vane turning shifted the first-stage stall-free performance and the knee in the stall-limit line to a lower engine speed. Opening the interstage bleed reduced the minimum rotor speed at which stall-free performance of the first stage was possible and tended to eliminate the knee in the stall-limit line.
Unsteady shock-induced separation of the ramp boundary layer was reduced and stabilized more effectively by external perforations than by external or internal slots. At Mach 2.0 peak total-pressure recovery was increased from 0.802 to 0.89 and stable mass-flow range was increased 185 percent over that for the solid ramp. Peak pressure recovery occurred just before instability. The 7 and one-third-diameter duct ahead of the engine reduced large total-pressure distortions but was not as successful for small distortions as obtained with throat bleed. By removing boundary-layer air the bypass nearly recovered the total-pressure loss due to the long duct.
Results of an experimental investigation of a symmetric crossing shock wave/turbulent boundary layer/bleed interaction are presented for a freestream unit Reynolds number of 1.68 x 10(exp 7)/m, a Mach number of 2.81, and deflection angles of 8 degrees. The data obtained in this study are bleed mass flow rate using a trace gas technique, qualitative information in the form of oil flow visualization, flow field Pitot pressures, and static pressure measurements using pressure sensitive paint. The main objective of this test is two-fold. First, this study is conducted to explore boundary layer control through mass flow removal near a large region of separated flow caused by the interaction of a double fin-induced shock wave and an incoming turbulent boundary layer. Also, a comprehensive data set is needed for computational fluid dynamics code validation.
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Explore the source record for details and available documents.
Explore the source record for details and available documents.
Explore the source record for details and available documents.
The transient performance of two concepts for control of vertical takeoff aircraft remote lift fans is analyzed and discussed. Both concepts employ flow transfer between pairs of lift fans located in separate parts of the aircraft in order to obtain attitude control moments for hover and low-speed flight. The results presented are from a digital computer, dynamic analysis of the YJ97/LF460 remote drive turbofan. The transient responses of the two systems are presented for step demands in lift and moment.