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
Boundary conditions for unsteady supersonic inlet analyses
New bleed and compressor face boundary conditions have been developed to improve the accuracy of unsteady supersonic inlet calculations. The new bleed boundary condition relates changes in the bleed hole discharge coefficient to changes in the local flow conditions; the local bleed flow rate can more than double as a shock moves forward over a bleed band in response to inlet flow disturbances. The stability margin of the inlet is strongly dependent on the throat bleed configuration since the locally rapid increase in bleed flow has a strong effect on the motion of the normal shock. The new compressor face boundary condition accounts for changes in the unsteady flow conditions at the compressor face by specifying the compressor face corrected mass flow or Mach number either as a constant or as a linear function of the stagnation conditions. The effects of inlet flow disturbances on the flow at the compressor face are represented more realistically with this new boundary condition than with traditional fixed static pressure or mass flow conditions. Euler calculations of the dynamic response of an inlet flow to a flow disturbance at the compressor face with 20- and 90-deg throat bleed hole angles are reported. These results indicate that an extra margin of stability for the inlet is obtained with 90-deg bleed holes because the increase in bleed flow rate as the shock moves forward over a bleed band is much larger for 90-deg holes than for 20-deg holes.