Showing papers on "Tip clearance published in 1981"

Stalling Pressure Rise Capability of Axial Flow Compressor Stages

Abstract: A procedure for estimating the maximum pressure rise potential of axial flow compressor stages is presented. A simplified stage average pitchline approach is employed so that the procedure can be used during a preliminary design effort before detailed radial distributions of blading geometry and fluid parameters are established. Semi-empirical correlations of low speed experimental data are presented that relate the stalling static-pressure-rise coefficient of a compressor stage to cascade passage geometry, tip clearance, bladerow axial spacing and Reynolds number. Blading aspect ratio is accounted for through its effect on normalized clearances, Reynolds number and wall boundary layer blockage. An unexpectedly strong effect of airfoil stagger and of the resulting flow coefficient of the stage’s vector triangle is observed in the experimental data. This is shown to be caused by the differing ability of different types of stage vector triangles to re-energize incoming low-momentum fluid. Use of a suitable “effective” dynamic head in the pressure rise coefficient gives a good correlation of this effect. Stalling pressure rise data from a wide range of both low speed and high speed compressor stages are shown to be in good agreement with these correlations.

On the Pressure Losses Due to the Tip Clearance of Centrifugal Blowers

Abstract: The pressure distribution along the shroud is measured for three types of centrifugal impeller at seven different values of tip clearance each. The change of input power due to a change of tip clearance is related to the effective blockage at the impeller tip. Since the change of input power is little for the test cases, the variation of local pressure gradient along the shroud is mostly attributed to the change of local pressure loss. The local pressure loss is related to the local tip clearance ratio and to the local pressure gradient based on the deceleration of relative velocity in the impeller. Since the local pressure gradient is largest near the throat of the impeller, for many impellers the clearance ratio at the throat is used as the representative value. The tip clearance loss is related to the clearance ratio and the pressure rise based on the deceleration of relative velocity in the impeller. A good correlation is observed in all cases at various flow rate.

Gas turbine engines

Abstract: A gas turbine engine blade tip seal consists of a sealing ring which is controlled by means of two annular control members The members are arranged such that one control member has a relatively rapid thermal response rate and the other control member has a relatively slow thermal response rate, the sealing ring being controlled such that a preferred tip clearance is maintained under varying engine operating conditions

Apparatus for minimizing and maintaining constant the blade tip clearance of axial-flow turbines in gas turbine engines

Abstract: Apparatus and method for minimizing and maintaining substantially constant the effective blade tip clearance between the outer free ends of rotor blades and an adjacent casing shroud of an axial-flow turbine of a gas turbine engine. The casing shroud includes a portion facing the hot gas stream and the outer rotor blade ends and a packing of high heat-resistance, high erosion-resistance ceramic elements secured to the shroud via a metal ring to face the blade tips. A heat insulator is interposed between the packing and the metal ring. A perforated conduit is mounted outside the metal ring for blowing cooling fluid against the metal ring on the side thereof remote from the insulator for equalizing the expansion of the metal ring to that of the turbine wheel. The ceramic packing includes projections facing the outer ends of the blades which are tapered so that they are easily broken off when coming into contact with the blades during running-in of the apparatus.

Laser-Optical Blade Tip Clearance Measurement System

Abstract: The need for blade tip clearance instrumentation has been intensified recently by advances in technology of gas turbine engines. A new laser-optical measurement system has been developed to measure single blade tip clearances and average blade tip clearances between a rotor and its gas path seal in rotating component rigs and complete engines. The system is applicable to fan, compressor and turbine blade tip clearance measurements. The engine mounted probe is particularly suitable for operation in the extreme turbine environment. The measurement system consists of an optical subsystem, an electronic subsystem and a computing and graphic terminal. Bench tests and environmental tests were conducted to confirm operation at temperatures, pressures, and vibration levels typically encountered in an operating gas turbine engine.

JT9D performance deterioration results from a simulated aerodynamic load test

Abstract: The results of testing to identify the effects of simulated aerodynamic flight loads on JT9D engine performance are presented. The test results were also used to refine previous analytical studies on the impact of aerodynamic flight loads on performance losses. To accomplish these objectives, a JT9D-7AH engine was assembled with average production clearances and new seals as well as extensive instrumentation to monitor engine performance, case temperatures, and blade tip clearance changes. A special loading device was designed and constructed to permit application of known moments and shear forces to the engine by the use of cables placed around the flight inlet. The test was conducted in the Pratt & Whitney Aircraft X-Ray Test Facility to permit the use of X-ray techniques in conjunction with laser blade tip proximity probes to monitor important engine clearance changes. Upon completion of the test program, the test engine was disassembled, and the condition of gas path parts and final clearances were documented. The test results indicate that the engine lost 1.1 percent in thrust specific fuel consumption (TSFC), as measured under sea level static conditions, due to increased operating clearances caused by simulated flight loads. This compares with 0.9 percent predicted by the analytical model and previous study efforts.

Analysis, design, fabrication and testing of an optical tip clearance sensor

Abstract: Analyses and the design, fabrication, and testing of an optical tip clearance sensor with intended application in aircraft propulsion control systems are reported. The design of a sensor test rig, evaluation of optical sensor components at elevated temperatures, sensor design principles, sensor test results at room temperature, and estimations of sensor accuracy at temperatures of an aircraft engine environment are discussed. Room temperature testing indicated possible measurement accuracies of less than 12.7 microns (0.5 mils). Ways to improve performance at engine operating temperatures are recommended. The potential of this tip clearance sensor is assessed.

Multistage Axial Compressor Program on Tip Clearance Effects.

Abstract: : Tip clearance has long been known to be a source of losses in axial compressors with cantilevered blades. The reasons for the losses, however, are not well understood and current practice in engine design still requires extensive effort to maintain constant minimal operating clearances over a wide range of conditions. The emphasis on clearance control may be appreciated by the typical observation that a ten percent change in peak static pressure rise in a compressor stage may occur for a fifty percent change in clearance. Clearances are typically in the one to five percent of major passage dimension range, and thus a small change in passage dimensions represents a large change in clearance. It is clear that, in general, it would be desirable that blading performance be less sensitive to changes in clearance. Less sensitivity would allow a general relaxation of the mechanical tolerances on a compressor assembly and provide more consistent transient performance. The aerodynamics of achieving such a situation are a challenge as the underlying requirement is improved performance at larger clearances. Work toward understanding the basic mechanisms of tip clearance effects with an emphasis on designing for clearance has been commenced at the Naval Postgraduate School turbopropulsion Laboratory (NPS/TPL). This report summarizes the preliminary work on the Multistage Compressor (MSC) facility at the Laboratory.

Progress report - multistage axial compressor program on tip clearance effects

Abstract: Abstract : Tip clearance has long been known to be a source of losses in axial compressors with cantilevered blades. The reasons for the losses, however, are not well understood and current practice in engine design still requires extensive effort to maintain constant minimal operating clearances over a wide range of conditions. The emphasis on clearance control may be appreciated by the typical observation that a ten percent change in peak static pressure rise in a compressor stage may occur for a fifty percent change in clearance. Clearances are typically in the one to five percent of major passage dimension range, and thus a small change in passage dimensions represents a large change in clearance. It is clear that, in general, it would be desirable that blading performance be less sensitive to changes in clearance. Less sensitivity would allow a general relaxation of the mechanical tolerances on a compressor assembly and provide more consistent transient performance. The aerodynamics of achieving such a situation are a challenge as the underlying requirement is improved performance at larger clearances. Work toward understanding the basic mechanisms of tip clearance effects with an emphasis on designing for clearance has been commenced at the Naval Postgraduate School turbopropulsion Laboratory (NPS/TPL). This report summarizes the preliminary work on the Multistage Compressor (MSC) facility at the Laboratory.