Development of an Experimental Data Base to Validate Compressor-Face Boundary Conditions Used in Unsteady Inlet Flow Computations
Dr. MiNos SajbenDr. Donald D. FreundDepartment of Aerospace Engineering and Engineering MechanicsUniversity of CincinnatiFinal Technical ReportResearch Grant Award fromNational Aeronautics and Space AdministrationNASA Lewis Research CenterPenod of Performance: January 1, 1997 to December 31, 1997May 15, 1998
Reflection of Large Amplitude Acoustic Pulses from an Axial Flow Compressor
The reliability of unsteady inlet flow computations may be seriously degraded by the lack of experimentally validated compressor-face (outflow) boundary conditions. It is common practice to require some flow variable to be constant (pressure, velocity or Mach number, etc.), but there is little if any documented evidence to support these assumptions. In addition, these assumptions do not have a basis in observation nor are they likely to actually occur during a rapid transient of a real inlet/engine system. The present paper presents measurements of acoustic reflection coefficients for an operating multistage compressor, a quantity appropriate for the characterization of the compressor face for computational purposes. The experiment used an impulse method, in which shortduration, large amplitude acoustic pulses (one msec, with a peak value of nearly 4% of the mean static pressure, respectively) were generated in an constantarea, annular inlet. The pulse and its reflection from the compressor-face were tracked by fast-response pressure transducers. Frequency-domain analysis of the data yields transfer functions that may be thought of as frequency-resolved reflection coefficients. It was demonstrated that transfer functions can be used for the accurate prediction of transients induced by arbitrary acoustic incident disturbances. None of the currently imposed compressor-face boundary conditions predict the data obtained in this study, indicating that current practices concerning outflow boundary conditions are in need of revision. Nomenclature
Dynamic Digital Blade Row Compression Component Stability Model. Model Validation and Analysis of Planar Pressure Pulse Generation and Two-Stage Fan Test Data.
Abstract : The objectives of this study were to determine the range of validity of a Dynamic Digital Blade Row Compression Component Stability Computer Model by comparing the results of model simulations with previously obtained test data and to determine the manner in which a fan component responded to imposed unsteady aerodynamics by examining calculated parameters. Validation of the modeling technique was obtained by simulating a facility and test conditions where the response of a two-stage fan to planar waves was determined. The simulation of the two-stage fan component was carried out using a quasi one-dimensional, pitch line, blade-row by blade-row model which used blade-row loss coefficients and deviation angles determined from analysis of clean-inlet-flow test data. This model was ideally suited for studying the fan response to the planar waves produced by the Planar Pressure Pulse Generator, a device which produced highly planar waves with peak-to-peak fluctuations of 10-30 percent of the mean total pressure over a frequency range of 42 to 800 Hz. The results of the simulations showed that the model accurately reproduced the waveshape, amplitude, and phase of the experimentally determined fan inlet and exit pressures and the fan inlet flow up to 75 Hz where the need for including unsteady blade circulation effects on blade characteristics began to become apparent. Further, the blade row model showed accurate simulation of the pressures in the interconnecting ducts by reproducing the wave and its harmonics up to approximately 118 Hz where the length of the volumes becomes too large to provide proper spatial definition of the harmonics of the standing waves although the fundamental frequency is properly reproduced.