ALBERT

Description: A linear solution modelling the shock-cell structure of an axisymmetric supersonic jet operated at slightly off-design conditions is developed by the method of multiple scales. The model solution takes into account the gradual spatial change of the mean flow in the downstream direction. Turbulence in the mixing layer of the jet has the tendency to smooth out the sharp velocity and density gradients induced by the shocks. To simulate this effect, eddy-viscosity terms are incorporated in the model. Extensive comparisons between the numerical results of the present model and experimental measurements gathered at the NASA Langley Research Center over the Mach number range of| Mj 2— Md 2|1.0 for underexpanded and overexpanded supersonic jets are carried out. Here Mj is the fully expanded jet Mach number and Mdis the design Mach number of the convergent-divergent nozzle. Very favourable agreement is found. This is especially true for the gross features of the shock cells, including the shock-cell spacings and the pressure amplitudes associated with the shocks. The measured data show that the pressure distributions over the first three or four shock cells usually are rich in fine structures. These fine structures are reproduced by the calculated results. Beyond the first few shock cells the model predicts that the shock-cell structure can be represented by a single Fourier mode of the mean flow. This is confirmed by a careful examination of the experimental data. The appropriate turbulent Reynolds number for shock-cell structure calculation is investigated. It is shown that the best choice is the same as the value found to give the best results for jet mean-flow calculation. The present model is used to explain some of the observed characteristics of broadband shock-associated noise. © 1985, Cambridge University Press. All rights reserved.

Description: The acoustic emission and initial shear layer of a cold jet issuing from an underexpanded sonic nozzle were measured where the lip thickness of the nozzle exit varied from 0.015 to 0.625 nozzle diameters. Because the amplitude of plume resonance (or screech) has been shown to be very sensitive to the initial conditions of the plume, the effects of nozzle lip thickness on this sound component in particular was investigated (although it will be shown that other noise components are also affected). Nearfield acoustic measurements have revealed the amplitude and frequency of certain modes of screech to be dependent on nozzle lip thickness. Fluctuating pressure measurements made on the nozzle exit surface have acoustic amplitudes in excess of near field microphone measurements. The dominant mode of instability that exists in the shear layer is also related to this geometric parameter. Detailed shear layer measurements were performed in an attempt to quantify any associated changes in the momentum thickness caused by altering the nozzle exit.

Description: Aircraft dynamic loads produced by engine exhaust plumes are examined for a class of military fighter and bomber configurations in model and full scale. The configurations examined are associated with the USAF F-15 and B-1B aircraft, and the US F-18 HARV and ASTOVL programs. The experience gained as a result of these studies is used to formulate a level of understanding concerning this phenomena that could be useful at the preliminary stage of propulsion/airframe design.

Description: SCIPVIS, the computational model discussed by Dash et al. (1985), is assessed in predicting the complicated flow structure associated with shock-containing plumes. In addition, the analysis in this study examines this code's applicability as a basic part of a program for estimating broadband shock noise radiation. The results of this study show that excellent agreement exists between predicted and measured static pressure distributions for both underexpanded and overexpanded flow cases considered. Of the three turbulence closure models incorporated in the SCIPVIS code, the kW model of Spalding produces the most uniform agreement with measurement. The k-epsilon-2 model of Launder consistently overestimates plume spreading for supersonic jets with exit Mach numbers in the 1-2 range. Dash's (1983) k-epsilon-2-cc, compressibility-corrected version of Launder's model underestimates plume spreading. Good qualitative agreement was also obtained between the measured longitudinal turbulence intensity and that predicted by the code for the same trial case. Comparison of measured and predicted broadband shock noise spectrum peak values were found to be in excellent agreement. This utilized a variant of the Harper-Bourne and Fisher (1973) phase-array model: the effective shock spacing was reinterpreted as the value of the end of the plume potential core, determined herein by the SCIPVIS code.

Description: The importance of reducing jet noise in both commercial and military aircraft applications has made jet acoustics a significant area of research. A technique for jet noise prediction commonly employed in practice is the MGB approach, based on the Lighthill acoustic analogy. This technique requires as aerodynamic input mean flow quantities and turbulence quantities like the kinetic energy and the dissipation. The purpose of the present paper is to assess existing capabilities for predicting these aerodynamic inputs. Two modern Navier-Stokes flow solvers, coupled with several modern turbulence models, are evaluated by comparison with experiment for their ability to predict mean flow properties in a supersonic jet plume. Potential weaknesses are identified for further investigation. Another comparison with similar intent is discussed by Barber et al. The ultimate goal of this research is to develop a reliable flow solver applicable to the low-noise, propulsion-efficient, nozzle exhaust systems being developed in NASA focused programs. These programs address a broad range of complex nozzle geometries operating in high temperature, compressible, flows. Seiner et al. previously discussed the jet configuration examined here. This convergent-divergent nozzle with an exit diameter of 3.6 inches was designed for an exhaust Mach number of 2.0 and a total temperature of 1680 F. The acoustic and aerodynamic data reported by Seiner et al. covered a range of jet total temperatures from 104 F to 2200 F at the fully-expanded nozzle pressure ratio. The aerodynamic data included centerline mean velocity and total temperature profiles. Computations were performed independently with two computational fluid dynamics (CFD) codes, ISAAC and PAB3D. Turbulence models employed include the k-epsilon model, the Gatski-Speziale algebraic-stress model and the Girimaji model, with and without the Sarkar compressibility correction. Centerline values of mean velocity and mean temperature are compared with experimental data.

Description: Previously cited in issue 24, p. 4250, Accession no. A81-49745

Description: Based on extensive work performed by Dr. Thomas H. Sobota (Advanced Projects Research Incorporated (APRI)) on swirling flows in circular-to-rectangular transition sections, a model assembly was designed and fabricated in support of a Phase 1 Small Business Innovation Research Contract between the NASA-Langley Research Center and APRI. This assembly was acoustically tested as part of this Phase 1 effort, the goal being to determine whether the controlled introduction of axial vorticity could affect the various noise generation mechanisms present in an underexpanded supersonic rectangular jet.