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In the present work, an existing transonic potential code is adapted to utilize the Multiple Level Adaptive technique proposed by A. Brandt. It is shown that order of magnitude improvements in speed and greatly improved accuracy over the unmodified code are achieved. Consideration is given to the difficulties of multi-grid programming, and possible future applications are surveyed.
The finite element codes were improved using Hermitian elements and numerical integration of element relations. The question of real variable versus complex variable formulation was resolved and an integrated civil engineering system was implemented on the Georgia Tech Cyber 70/74. Efforts are underway to restructure the program to obtain the most efficient use of array storage.
A single stage axial compressor with distorted inflow is studied. The inflow distortion occurs far upstream and may be a distortion in stagnation temperature, stagnation pressure or both. The blade rows are modeled as semi-actuator disks. Losses, quasi-steady deviation angles, and reference incidence correlations are included in the analysis. Both subsonic and transonic relative Mach numbers are considered. A parameter study is made to determine the influence of such variables as Mach number and swirl angle on the attenuation of the distortion.
The results of weakly viscous flow analysis are presented. The flow field, including molecular transport, is computed with the aid of a bicharacteristic method. The bow shock wave and the internal shock wave are computed with the aid of a three-dimensional shock wave fitting procedure. Characteristic equations are presented, and numerical integration procedure is discussed. Here, an inverse marching scheme is employed in which the solution is obtained on space-like planes of constant x and on space curves defined by the intersections of the internal shock wave with the solid boundaries. The distance between solution planes is arrived at by the Courant-Friedrichs-Lewy stability criterion.
The acoustic field within a duct containing a nonuniform steady flow was predicted. This analysis used the finite element method to calculate the velocity potential within the duct.
(Previously announced in STAR as N82-24165)
(Previously cited in issue 24, p. 4248, Accession no. A81-49736)
(Previously cited in issue 07, p. 963, Accession no. A82-19212)
Previously cited in issue 19, p. 3257, Accession no. A81-42186
A mathematical model was developed for spinning mode acoustic radiation from a thick wall duct without flow. This model is based on a series of experiments (with and without flow). A nearly pure azimuthal spinning mode was isolated and then reflection coefficients and far field pressure (amplitude and phase) were measured. In our model the governing boundary value problem for the Helmholtz equation is first converted into an integral equation for the unknown acoustic pressure over a disk, S1, near the mouth of the duct and over the exterior surface, S2, of the duct. Assuming a pure azimuthal mode excitation, the azimuthal dependence is integrated out which yields an integral equation over the generator C1 of S1 and the generator C2 of S2. The sound pressure on C1 was approximated by a truncated modal expansion of the interior acoustic pressure. Piecewise linear spline approximation on C2 was used.
A numerical procedure which solves the parabolized Navier-Stokes (PNS) equations on a body fitted mesh was used to compute the flow about the forebody of an advanced tactical supercruise fighter configuration in an effort to explore the use of a PNS method for design of supersonic cruise forebody geometries. Forebody flow fields were computed at Mach numbers of 1.5, 2.0, and 2.5, and at angles-of-attack of 0 deg, 4 deg, and 8 deg. at each Mach number. Computed results are presented at several body stations and include contour plots of Mach number, total pressure, upwash angle, sidewash angle and cross-plane velocity. The computational analysis procedure was found reliable for evaluating forebody flow fields of advanced aircraft configurations for flight conditions where the vortex shed from the wing leading edge is not a dominant flow phenomenon. Static pressure distributions and boundary layer profiles on the forebody and wing were surveyed in a wind tunnel test, and the analytical results are compared to the data. The current status of the parabolized flow flow field code is described along with desirable improvements in the code.
Previously cited in issue 17, p. 2675, Accession no. A82-35195
Full-scale, in-flight measurements of the boundary-layer thickness, velocity profile, and flow angle have been made at several sample collection stations on the fuselage of the NASA CV 990. These results are given as functions of Mach number, Reynolds number, yaw, and angle of attack.
Previously cited in issue 07, p. 965, Accession no. A82-19778
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One of the primary goals is to develop an approximate but sufficiently accurate analysis for the problem of electromagnetic (EM) plane wave scattering by an open ended, perfectly-conducting, semi-infinite hollow circular waveguide (or duct) with a thin, uniform layer of lossy or absorbing material on its inner wall, and with a simple termination inside. The less difficult but useful problem of the EM scattering by a two-dimensional (2-D), semi-infinite parallel plate waveguide with an impedance boundary condition on the inner walls was chosen initially for analysis. The impedance boundary condition in this problem serves to model a thin layer of lossy dielectric/ferrite coating on the otherwise perfectly-conducting interior waveguide walls. An approximate but efficient and accurate ray solution was obtained recently. That solution is presently being extended to the case of a moderately thick dielectric/ferrite coating on the walls so as to be valid for situations where the impedance boundary condition may not remain sufficiently accurate.
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