| Summary: | Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1999. === Includes bibliographical references (p. 195-199). === A transient testing technique for the study of jet noise was investigated and assessed. A shock tunnel facility was utilized to produce short duration, 10-20 millisecond, under expanded supersonic hot air jets from a series of scaled nozzles. The primary purpose of the facility is to investigate noise suppressor nozzle concepts relevant to supersonic civil transport aircraft applications. The shock tube has many strengths; it is mechanically simple, versatile, has low operating costs, and can generate fluid dynamic jet conditions that are comparable to aircraft gas turbine engine exhausts. Further, as a result of shock heating, the total temperature and pressure profiles at the nozzle inlet are uniform, eliminating the noise associated with entropy non-uniformities that are often present in steady state, vitiated air facilities. The primary drawback to transient testing is the brief duration of useful test time. Sufficient time must be allowed for the nozzle flow and free jet to reach a quasi-steady-state before acoustic measurements can be made. However, if this constraint is met, the short run times become advantageous. The test articles are only exposed to the high temperature flow for a fraction of a second, and can be constructed of relatively inexpensive stereo-lithography or cast aluminum. A comparison between shock tunnel transient noise data and steady-state data is presented to ascertain the usefulness of the technique to make acoustic measurements on scaled nozzles. Three types of nozzles are compared in the assessment effort: (1) a series of 0.64 - 1.9 cm exit diameter small-scale round nozzles that can be operated at transient and cold-flow steady-state conditions at the MIT facility for in-house comparison, (2) a series of 5.1 - 10.2 cm exit diameter ASME standard axisymmetric nozzles, and (3) a 1/1 2th scale version of a modern mixer-ejector nozzle. Scaled versions of nozzles (2) and (3) were tested at Boeing's steady-state low speed aeroacoustic facility for comparison to the transient shock tube noise data. The assessment establishes the uncertainty bounds on sound pressure level measurements over the range of frequency bands, nozzle pressure ratios (1.5 - 4.0), total temperature ratios (1.5 - 3.5), and nozzle scales for which the facility can be employed as a substitute and/or as a complimentary mode of investigation to steady-state hot-flow test facilities. Far-field narrowband spectra were obtained at directivity angles from 65 to 145 degrees and the data were extrapolated to full-scale flight conditions consistent with FAR-36 regulations. Nozzle pressure ratio and total temperature ratio were repeatable to within ± 1 percent of desired conditions. The constraint of short test duration is shown to be alleviated through the use of multiple runs to reduce the uncertainty associated with making transient acoustic measurements. Sound pressure level versus frequency trends with nozzle pressure ratio and directivity angle are shown to be comparable between the steady-state and transient noise data for all three nozzle types. The small scale nozzles exhibited agreement to within ± 1 - 2 dB over a full-scale frequency range of 50 - 1250 Hz. The ASME nozzle results demonstrated that the transient noise data replicates the Boeing steady-state data to within 2 - 3 dB on SPL versus full-scale frequency from 250 - 6300 Hz, as well as OASPL and PNL versus directivity angle. The magnitude of EPNL values are shown to agree to within 1 - 3 dB depending on test condition and nozzle scale. The mixer-ejector model exhibited agreement with the steady-state noise data to within 2 - 5 dB over a frequency range of 500 - 6300 Hz for all directivity angles. OASPL and PNL versus directivity angle noise data exhibited agreement with magnitude to within 1 - 4 dB. Steady-state trends with MAR, azimuthal angle, and EPNL were also present in the transient noise data. === by Daniel Robert Kirk. === S.M.
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