Showing papers on "Supersonic speed published in 1972"

On the noise of a nearly ideally expanded supersonic jet

Abstract: A noise generation mechanism for a nearly ideally expanded supersonic jet is proposed. It is suggested that the dominant part of the noise of a supersonic jet is generated at two rather localized regions of the jet. These regions are located at distances quite far downstream of the nozzle exit. Large-scale instabilities of the jet flow are believed to be responsible for transferring the kinetic energy of the jet into noise radiation. An analysis based on a simple mathematical model reveals that two large-scale unstable waves are preferentially amplified in a supersonic jet. The rapid growth of these waves causes the oscillations of the jet to penetrate the mixing layer at two locations and to interact strongly with the ambient fluid there. This gives rise to intense noise radiation. Theoretical results based on the proposed noise generation mechanism are found to compare favourably with experimental measurements. A simple scaling formula is also derived.

The effect of injection angle on the interaction between sonic secondary jets and a supersonic free stream

Abstract: An experimental investigation, as part of a research program on the development of technology for the design of supersonic combustion ramjets, has been conducted to determine the effect of injection angle on the jet penetration, mixing rate, and airstream total-pressure recovery downstream of five laterally spaced sonic hydrogen jets flush mounted on a flat plate. Results of this investigation indicated that at lower injection angles less free-stream momentum loss was required to turn and accelerate the injected gas downstream and, thereby, less flow disturbance and total-pressure loss were produced. In addition, lower injection angles resulted in improved fuel distribution and faster mixing of the injected gas with the free stream. A correlating parameter, developed from considerations of the effective-momentum-flux differences between the injected gas and the free-stream air, predicted greater penetration and faster mixing for the lower injection angles.

Material Phase Transformation Effects upon Performance of Spaced Bumper Systems

Abstract: 10 Carriere, P. and Sirieix, M., "Facteurs d'lnfluence du Receullement d'un Ecoulement Supersonique," Proceedings of the 10th International Congress of Applied Mechanics, Stresa, Italy, 1960. 11 McDonald, H., "An Analysis of the Turbulent Base Pressure Problem in Supersonic Axisymmetric Flow," The Aeronautical Quarterly, Vol. XVI, May 1965, pp. 97-121. 12 Korst, H. H., Chow, W. L. and Zumwalt, G. W., "Research on Transonic and Supersonic Flow of a Real Fluid at Abrupt Increases in Cross Section (With Special Consideration of Base Drag Problems)," ME TR 392-5, Oct. 1964, University of Illinois, Urbana, 111. 13 McDonald H. and Hughs, P. F., "A Correlation of High Subsonic Afterbody Drag in the Presence of a Propulsive Jet or Support Sting," Journal of Aircraft, Vol. 2, No. 3, May-June, 1965, pp. 202-207. Dixon, R. J. and Page, R. H., "Interdependence of Base Pressure and Base Heat Transfer," ARS Journal, Vol. 31, No. 12, Dec. 1961, pp. 1785-1786. 15 Collines, D. J., Lees, L. and Roshko, A., "Near Wake of a Hypersonic Blunt Body with Mass Additions," AIAA Journal, Vol. 8, No. 5, May 1970, pp. 833-842. 16 Mueller, T. J., United Aircraft Research Laboratory, East Hartford, Conn. 1965, unpublished data.

Noise field of a supersonic Mach 1.5 cold model jet.

Abstract: Systematic surveys of both the near and far noise fields were made for a supersonic (nozzle design Mach number 1.5) cold model jet. The purpose of this investigation was to extend the existing understanding of the noise field of a supersonic shock-free jet at a moderate exit Mach number and the variations in its acoustic behavior when the nozzle was operated at its underexpanded mode, where strong shock waves were present. It was found that the broad-band pressure spectra of apparent sources in each of the individual characteristic flow regions were grossly similar. The strongest source for the shock-free jet was located in the region between the laminar core tip and the supersonic core tip, whereas for the underexpanded jet the strongest source was found near the middle of the flow region containing repetitive shock waves. It is shown that the significant increase in the noise output at the underexpanded mode of operation was primarily due to high-frequency components.

Seismic source time function of propagating longitudinal‐shear cracks

Abstract: An expression of stress components corresponding to a time-dependent dislocation (discontinuity of displacement) is obtained for the case of a two-dimensional longitudinal-shear crack, on the basis of the Green function representation theorem of the elastic-wave equation. This expression is used to examine various dislocations from the viewpoint of the boundary condition that should be satisfied on the fault plane. It is shown that the dislocation functions consistent with propagating cracks have quite different characteristics between the supersonic propagation with higher than shear-wave velocity and the subsonic propagation with lower than shear-wave velocity. For supersonic propagation, the source time function giving correct rupture propagation represents a step-function type of change in the particle velocity, whereas the propagating cracks have a singularity similar to the static cracks at the edge in the case of subsonic propagation. The source time function proposed by J. N. Brune is applicable to supersonic propagation if a suitable correction is made, but this time function does not represent the correct behavior of the cracks propagating at a subsonic velocity. A source time function that is applicable to subsonic rupture propagation is also presented in this paper.

Three-dimensional, minimum-time turns for a supersonic aircraft

Abstract: Altitude, bank angle and thrust program for minimizing time required by supersonic aircraft to turn through specified heading angle and reach required energy

Reduction of wave drag by antisymmetric arrangement of wings and bodies.

Abstract: The wave interference effects for bodies or wings in a mirror-symmetric arrangement, and in an antisymmetric arrangement are discussed. It is shown that while in the case of a mirror-symmetric arrangement large adverse interference effects can be observed, antisymmetric arrangements provide comparatively much smaller wave drags. The single continuous wing panels also adapt themselves more readily to varying angles of obliquity, and hence, to varying flight speeds. A detailed review is presented of the previous work on the aerodynamic properties and flight stability of oblique elliptic wing combinations. A possible mode of application of these combinations to transport aircraft operating at moderate supersonic speeds is suggested.

Swirling Base Injection for Supersonic Combustion Ramjets

Abstract: Experimental evaluation of the swirling base injection proposed by Swithenbank and Chigier (1969) for application in supersonic combustion ramjets or scramjets. This concept of accelerated mixing in supersonic streams through swirl was tested, but the results indicate that swirl does not produce any enhancement of mixing.

Removal of particulate matter with supersonic droplets

Abstract: Water droplets traveling at supersonic velocities are utilized to remove particulate matter from a gas stream in which the matter is entrained. The gas stream, such as an industrial effluent containing particulate contaminants, flows through a duct at a relatively low velocity. The outlet of a convergingdiverging nozzle communicates with the duct at a mixing region thereof, and steam under pressure at subsonic velocity is supplied to the inlet of the nozzle. The steam reaches sonic velocity in the throat of the nozzle, is accelerated to supersonic velocity in the diverging section of the nozzle, and undergoes isentropic (or essentially isentropic) expansion to cause partial condensation, resulting in the formation of droplets of a size on the order of 0.1 micron. These supersonic droplets are injected into the duct and combine with the contaminants to thereby increase the size of particles carried by the gas stream downstream from the mixing region. A cyclone separator is located downstream to separate the particles of increased size from the gas stream.

Prediction of the Spectral Content of Combination Tone Noise

Abstract: Combination tone noise is generated by a pattern of rotating shock waves in the inlet of turbofan engines when the relative tip speed of the fan blades is supersonic. A method is presented for determining the expected distribution of sound power among the harmonics of engine rotation frequency, based on the spectral analysis of an almost periodic succession of pulses. The spectral distribution of combination tone noise is found to depend on two statistical — crfl, the standard deviations of the sequence of shock wave amplitudes; and cr£, the standard deviation of the sequence of time intervals between successive shock waves. The spectral distribution of sound power is found to depend more critically on

Jet-interaction-induced separation - The two-dimensional problem.

Abstract: An experimental study supporting the development of an analytical model for jet-interaction-induced separation of supersonic turbulent boundary layers is discussed. Extensive flat-plate tests were conducted at a Mach number of four and two Reynolds numbers. Surface pressures were recorded fore and aft of the four, sonic, normal jet slots tested. Shadowgraphs, taken through glass-ported side plates, were made of both the interacting and freejet plume characteristics. Generalized correlations showed that the entire problem scales directly with the observed shock heights and that these shock heights are predictable from freejet considerations.

A method of through computation for two- and three-dimensional supersonic flows. I

Abstract: THE first part develops and applies a method of through computation to two- and three-dimensional supersonic stationary flows of an ideal (non-viscous and thermally non-conducting) gas The method is explicit and ensures fairly weak “smearing” of the shock waves Its scope is illustrated by computations of supersonic flow in rotationally symmetric nozzles and jets

Sonic‐Boom Propagation through a Stratified Atmosphere

Abstract: The evolution of the sonic‐boom signature from a supersonic aircraft is discussed and treated. This treatment is the basis for a general algorithm and computer program for sonic‐boom prediction in a stratified atmosphere with steady winds. The treatment is based upon geometric acoustics with calculation of nonlinear distortion, and takes arbitrary maneuvers of the aircraft into account.

Pressure-fluctuation inputs and response of panels underlying attached and separated supersonic turbulent boundary layers

Abstract: Results are presented for an investigation of surface pressure fluctuations and response of panels underlying attached and separated turbulent boundary layers and shock waves. Tests conducted at transonic and supersonic Mach numbers to 3.6 to study the pressure fields. Assorted fixed edge flat panels were tested at Mach numbers from 1.6 to 3.6 in attached and completely separated flow fields, and also in mixed flow with a step-induced shock wave oscillating on the panels. The surface pressure fluctuations are described in terms of broadband rms, spectral density, and spatial correlation information. The effectiveness of parameters for scaling the pressure fluctuations is also illustrated. Measurements of the amplitude and strain response of the panels are compared with response computations by the normal mode method of analysis.

Aeroelastic Optimization of a Panel in High Mach Number Supersonic Flow

Abstract: Solution for a least-weight skin thickness distribution for a panel with a flutter parameter constraint. This panel weighs less than any similar constant thickness panel, but has the same critical supersonic panel flutter parameter. The panel rests on simple supports and is of sandwich construction. The span to chord ratio is large enough that the inertial, elastic, and aerodynamic behavior is one-dimensional. The Mach number is great enough that the aerodynamic forces acting on the upper panel surface may be accurately described by quasi-steady, linearized, supersonic aerodynamic theory. The final optimum design is obtained from theoretical and numerical methods adapted from optimal control theory. The results of this investigation show that the optimal panel thickness distribution is symmetric about the panel chord midpoint. Compared to a reference panel with constant thickness, optimum panels are found to be nearly 12% lighter.

Impact Probe Displacement in a Supersonic Turbulent Boundary Layer

Abstract: Eight circular pitot probes ranging in size from 2 to 70 percent of the boundary-layer thickness were tested to provide experimental probe displacement results in a two-dimensional turbulent boundary layer at a nominal free-stream Mach number of 2 and unit Reynolds number of 8 million per meter The displacement obtained in the study was larger than that reported by previous investigators in either an incompressible turbulent boundary layer or a supersonic laminar boundary layer The large probes indicated distorted Mach number profiles, probably due to separation When the probes were small enough to cause no appreciable distortion, the displacement was constant over most of the boundary layer The displacement in the near-wall region decreased to negative displacement in some cases This near-wall region was found to extend to about one probe diameter from the test surface

Dual purpose circulation control airfoil

Abstract: A dual purpose circulation control airfoil is designed for operation at high speeds in the vicinity of transonic speed and for low speed applications such as landing speeds. A blunt trailing edge possesses a varying radius of curvature to comprise a high speed and low speed coanda surface. For high speed operation, air is blown from a slot over the high speed surface, moving upper surface stagnation point further aft and increasing lift. At low speed operation air is blown over the low speed surface to produce high lift for low landing speeds. In addition, a surface forward of the blunt edge is provided for supersonic expansion in transonic flight.

Theory of Exhaust-Plume/Boundary-Layer Interactions at Supersonic Speeds

Abstract: At supersonic speeds, an under-expande d rocket exhaust jet produces a deflection of the external flow and generates a rise in pressure that communicates upstream through the surface boundary layer For sufficiently large pressure ratios, the viscous layer separates some distance forward of the base and may seriously affect vehicle stability The present investigation develops an integral method for predicting some of the main features of the nearfield interaction between the exhaust plume and the surface boundary layer The analysis is not restricted to laminar flow and extends to turbulent interactions by utilizing a simplified eddy-diffusivity model The method also applies to the interaction that occurs in the wake of inclined or asymmetric bodies, and results for the flow over a flat plate at angle of attack are given

Analysis of Supersonic-Hypersonic Flutter of Lifting Surfaces at Angle of Attack

Abstract: Several important features of supersonic-hypersonic flutter of lifting surfaces at angle of attack are highlighted in an exploratory study. Three simple analytical methods—a modified strip analysis incorporating shock expansion theory, modified Newtonian flow theory, and local flow piston theory—are employed in order to avoid the usual limitations of piston theory and Newtonian flow theory with regard to Mach number range, angle-of-attack range, and airfoil section. Illustrative flutter calculations have been made for two rectangular wings with diamond airfoils and with pitch and translational degrees of freedom. Results from the modified strip analysis and from the modified Newtonian flow theory are in reasonable agreement with limited available experimental data for Mach number 10, but results from the local flow piston theory are unconservative. The calculations show a typical degradation of flutter-speed index with increasing angle of attack, which at low angles of attack is most pronounced for thin sections. For some airfoil shapes, however, a forward location of the center of gravity may mitigate the degradation at low to moderate angles of attack and essentially postpone it until shock detachment conditions are approached. As angle of attack is increased, or Mach number is decreased toward the shock detachment condition, a sharp drop in flutter speed index is predicted by the methods that account for detachment. Thus, the vicinity of shock detachment is indicated to be a critical region for supersonic-hypersonic flutter.

Liquid jet injection into a supersonic flow

Abstract: T study of a liquid jet injected normal to a supersonic gas crossflow is one of significant practical import and stimulating theoretical interest. There has been, however, a relative scarcity of experimental data pertaining to many details of the process, particularly the primary decomposition mechanism. The immediate stimulus and sources of the initial direction for this study were the experiments of Sherman and Schetz, wherein the presence on the jet of large axial waves, which appeared to create the dominant means of jet decomposition by gross detachment of large masses, was noted.

Dimer Concentrations in Supersonic Free Jets

Abstract: Measurements have been made of the concentration of dimers produced during the expansion from a supersonic nozzle. The gases studied were: Ar, N2,O2, CO, CO2, C2H4, and N2O. It is shown that an effective specific heat ratio of the expanding gas offers a means of correlating the dimer concentration in terms of the source gas pressure and temperature. The two nozzles employed had opening diameters of 0.14 mm and 0.065 mm. An attempt to obtain a universal dependence for all gases in terms of reduced source parameters was partially successful. Source temperature was varied between 90 °K and 295 °K.

Influence of Contraction Section Shape and Inlet Flow Direction on Supersonic Nozzle Flow and Performance

Abstract: Wall static pressure measurements and performance parameters are presented for axisymmetric supersonic nozzles with relatively steep convergent sections and comparatively small radius-of-curvature throats. The nozzle walls were essentially adiabatic. These results are compared with those obtained in other nozzles tested previously to appraise the influence of contraction shape on performance. Both the flow coefficient and the thrust were less than the corresponding values for one-dimensional, isentropic, plane flow for both the axial and radial inflow nozzles considered, but the specific impulse, the most important performance parameter, was found to be relatively unchanged. The thrust decrement for the axial inflow nozzles was established primarily by the shape of the contraction section, and could be estimated reasonably well from a conical sink flow consideration. The radial inflow nozzle has a potential advantage from a cooling point of view if used in a rocket engine.

Body Alone Aerodynamics of Guided and Unguided Projectiles at Subsonic, Transonic and Supersonic Mach Numbers

Abstract: : Several theoretical and empirical methods are combined into a single computer program to predict lift, drag, and center of pressure on bodies of revolution at subsonic, transonic, and supersonic Mach numbers. The body geometries can be quite general in that pointed, spherically blunt, or truncated noses are allowed as well as discontinuities in nose shape. Particular emphasis is placed on methods which yield accuracies of ninety percent or better for most configurations but yet are computationally fast. Theoretical and experimental results are presented for several projectiles and a computer program listing is included as an appendix.