Showing papers on "Supersonic speed published in 1974"

Centerline Velocity Decay of Compressible Free Jets

Abstract: Theme A empirically determined formulation for the eddy viscosity is introduced into the turbulent axially symmetric compressible free jet theory of Kleinstein. * The easily evaluated algebraic equation that results shows excellent agreement with an extensive compilation of experimental data for the centerline velocity decay in the main region. It is found that the eddy viscosity can be treated as constant, with functional dependence on only the exit Mach number and the jet-to-freestream density ratio. The theory presented is applicable to the main region of most classes of free jets, including heterogeneous, nonisothermal, and subsonic and properly-expanded supersonic flows.

Effect of the Adverse Pressure Gradient on Vortex Breakdown

Abstract: of the Influence of the Turbulent Boundary Layer on the Pressure Distribution over a Rigid Two-Dimensional Wavy Wall," TN D-6477, Aug. 1971, NASA. 4 Lighthill, M. J., "On Boundary Layers and Upstream Influence II. Supersonic Flows without Separation," Proceedings of the Royal Society, Vol. A217, 1953, pp. 478 and 504; see also Quarterly Journal of Mechanics, Vol. 3, 1950, p. 303. 5 Benjamin, T. B., "Shearing Flow over a Wavy Boundary," Journal of Fluid Mechanics, Vol. 6, 1959, p. 161. 6 Miles, J. W., "On Panel Flutter in the Presence of a Boundary Layer," Journal of Aerospace Sciences, Vol. 26, No. 2, Feb. 1959, pp. 81-93. 7 McClure, J. D., "On Perturbed Boundary Layer Flows," Rept. 62-2, June 1962, M.I.T. Fluid Dynamic Research Lab., Cambridge, Mass. 8 Anderson, W. J. and Fung, Y. C, "The Effect of an Idealized Boundary Layer on the Flutter of Cylindrical Shells in Supersonic Flow," GALCIT Structural Dynamics Rept. SM62-49, Dec. 1962, Graduate Aeronautical Lab., California Institute of Technology, Pasadena, Calif. 9 Zeydel, E. F. E., "Study of the Pressure Distribution on Oscillating Panels in Low Supersonic Flow with Turbulent Boundary Layer," NASA CR-691, Feb. 1967, Georgia Institute of Technology, Atlanta, Ga. 10 Do well, E. H., "Generalized Aerodynamic Forces on a Flexible Plate Undergoing Transient Motion in a Shear Flow with an Application to Panel Flutter," AIAA Journal, Vol. 9, No. 5, May 1971, pp. 834-841. 11 Ventres, C. S., "Transient Panel Motion in a Shear Flow," AMS Rept. 1062, Aug. 1972, Princeton Univ., Princeton, N.J. 12 Garrick, I. E. and Rubinow, S. I., "Theoretical Study of Air Forces on an Oscillating or Steady Thin Wing in a Supersonic Main Stream," TN 1383, July 1947, NACA. 13 Yates, J. E., "A Study of Panel Flutter with the Exact Method of Zeydel," NASA CR-1721, Dec. 1970, Aeronautical Research Associates of Princeton, Princeton, N.J. 14 Do well, E. H. and Ventres, C. S., "Derivation of Aerodynamic Kernel Functions," AIAA Journal, Vol. 11, No. 11, Nov. 1973, pp. 1586-1588.

The near wall jet of a normally impinging, uniform, axisymmetric, supersonic jet

Abstract: The near wall jet produced by directing a uniform axisymmetric jet of air normally onto a large flat plate has been investigated experimentally and theoretically for four jets in the Mach number range 1·64–2·77. Detailed measurements of the surface pressure and shadowgraph and surface flow pictures are presented. The results show that the mechanism which mainly determines the supersonic near wall jet is the jet-edge expansion and its reflexions from the sonic line and the wall-jet boundaries. The near wall jet is found to consist of an alternating series of expansion and recompression regions whose strengths depend on the jet Mach number and decay with distance. At Mach numbers of 2·4 and above, shock waves are observed in the first recompression region and at a Mach number of 2·77 the boundary layer separates locally. Further out, viscous effects become increasingly important and a constant-pressure shear flow is established at a distance which increases with jet Mach number. The application of the method of characteristics in an approximate manner reproduces a number of the features of the near wall jet which are observed experimentally.Pressure distributions obtained in the shock layer show that a stagnation bubble can occur and that its occurrence depends on factors such as the flow upstream of the nozzle. The wall-jet region is found to be largely independent of whether or not a bubble occurs in the shock layer.

Numerical methods for the design and analysis of wings at supersonic speeds

Abstract: Numerical methods for the design and analysis of arbitrary-planform wings at supersonic speeds are reviewed. Certain deficiencies are revealed, particularly in application to wings with slightly subsonic leading edges. Recently devised numerical techniques which overcome the major part of these deficiencies are presented. The original development as well as the more recent revisions are subjected to a thorough review.

A finite difference method for the solution of the transonic flow around harmonically oscillating wings

Abstract: A finite difference method for the solution of the transonic flow about a harmonically oscillating wing is presented. The partial differential equation for the unsteady transonic flow was linearized by dividing the flow into separate steady and unsteady perturbation velocity potentials and by assuming small amplitudes of harmonic oscillation. The resulting linear differential equation is of mixed type, being elliptic or hyperbolic whereever the steady flow equation is elliptic or hyperbolic. Central differences were used for all derivatives except at supersonic points where backward differencing was used for the streamwise direction. Detailed formulas and procedures are described in sufficient detail for programming on high speed computers. To test the method, the problem of the oscillating flap on a NACA 64A006 airfoil was programmed. The numerical procedure was found to be stable and convergent even in regions of local supersonic flow with shocks.

Oblique-wing supersonic aircraft

Abstract: An aircraft including a single fuselage having a main wing and a horizontal stabilizer airfoil pivotally attached at their centers to the fuselage. The pivotal attachments allow the airfoils to be yawed relative to the fuselage for high speed flight, and to be positioned at right angles with respect to the fuselage during take-off, landing, and low speed flight. The main wing and the horizontal stabilizer are upwardly curved from their center pivotal connections towards their ends to form curvilinear dihedrals.

Method of and apparatus for preventing leading edge shocks and shock-related noise in transonic and supersonic rotor blades and the like

Abstract: This disclosure is concerned with novel rotor blade and similar foil designs that, by critical skewing of intermediate blade regions where the airflow is supersonic, prevents leading edge shocks and shock-related noise.

Laser-Doppler velocity measurements in an under-expanded free jet

Abstract: Velocity measurements in an under-expanded supersonic free jet have been made using the Fabry-Perot laser-Doppler technique. The jet was produced by a nozzle of exit diameter 27 mm operating at a pressure ratio P0/Pinfinity=66. Measurement of the complex structure, characteristic of this type of jet, on such a small scale proved to be an excellent test of the capability of the Fabry-Perot laser-Doppler technique for making air velocity measurements.Sufficient scattering particles were present in the flow to make artificial seeding unnecessary. The particles were found to respond to the velocity change across the Mach disc (a normal shock) in a distance of similar 01 mm. Laser-Doppler velocity measurements were compared with velocity measurements deduced from Pitot pressure measurements. The difference in results between the two methods was < 1%. Sufficient laser-Doppler velocity measurements were obtained to construct the axial and four transverse velocity profiles which clearly showed the positions and magnitudes of the major features characteristic of an under-expanded free jet.

A Simplified Dynamic Model of Parachute Inflation

Abstract: This paper describes a dynamic inflation model for parachutes which predicts increased dimensionless inflation times and increased dimensionless inflation forces observed at high altitudes. As altitude is increased, greater relative parachute inertia results in increased inflation times, and greater relative total system inertia results in increased maximum inflation forces. The effect of Mach number on inflation force is also predicted by the inflation model.

Supersonic centrifugal compressors

Abstract: A supersonic centrifugal compressor comprises a rotor located in a housing having a fluid intake eye. The fluid (air for instance) successively travels through an intake region wherein the rotor has a small number of blades which deflect the fluid tangentially by a small amount only, then through a compression region wherein the rotor has a higher number of blades producing tangential and meridian flow deflection. Last, the fluid flows substantially radially with respect to the rotor into a stationary diffuser.

New Method for Supersonic Boundary-Layer Separations

Abstract: An efficient numerical solution algorithm is presented for solving the interacting supersonic laminar boundarylayer problem. The method employs a time dependent approach with the alternating direction implicit (ADI) scheme and directly accounts for the necessary downstream boundary condition. Solutions are presented for MOO = 3 cold wall boundary-layer flow over a family of compression ramps with regions of reverse flow. Good comparison is given with experimental data and Navier-Stokes solutions for M^ =4 and 6, adiabatic wall, separated flow up a 10° compression ramp.

Supersonic Flow in the Corner Formed by Two Intersecting Wedges

Abstract: Theme E12 for over a decade, have investigated the supersonic flowfield in the axial corner formed by two intersecting wedges. The understanding and prediction of this type of interference flow is of considerable interest to the designer of high-performance hypersonic aircraft and spacecraft (Space Shuttle Orbiter) because of the high local heating that occurs in this region. Such flows are encountered, for example, at wing-body junctures, at control surface roots, and within supersonic inlets. Theoretical attempts at modeling the corner flow problem have been made in the past,' but these employed simplifying or incorrect assumptions and thus resulted in questionable solutions. According to the recent survey paper of Korkegi, "there exists no adequate method of predicting even the inviscid flow structure." In the present paper, a numerical solution for the inviscid corner flow is computed using a second-order accurate shock-capturing technique.

Oscillatory supersonic kernel function method for interfering surfaces

Abstract: In the method presented in this paper, a collocation technique is used with the nonplanar supersonic kernel function to solve multiple lifting surface problems with interference in steady or oscillatory flow. The pressure functions used are based on conical flow theory solutions and provide faster solution convergence than is possible with conventional functions. In the application of the nonplanar supersonic kernel function, an improper integral of a 3/2 power singularity along the Mach hyperbola is described and treated. The method is compared with other theories and experiment for two wing-tail configurations in steady and oscillatory flow.

Design and Operation of Aircraft to Minimize Their Sonic Boom

Abstract: Means of reducing or eliminating the sonic boom through aerodynamic design or aircraft operation are discussed. These include designing aircraft to minimize or eliminate certain features of the overpressure signature, operating aircraft at slightly supersonic speeds so that the sonic boom does not reach the ground, and seeking reductions through the high-altitude, high-speed flight conditions of hypersonic transports. A simple computer program has been developed that determines the area development of the equivalent body of revolution required to minimize various sonic boom signature parameters.

Supersonic shock wave compressor diffuser with circular arc channels

Abstract: A supersonic shock wave compressor diffuser forms a concentric annulus about a radial compressor having a central axis. Circular channels diverge with an increasing divergence angle as they extend along an arcuate longitudinal center line from an inner circumference very near the periphery of the compressor to the outer circumference of the diffuser. Shock waves may occur within the channels near the inner circumference of the diffuser or may occur within a vaneless diffusion space adjacent the periphery of the compressor and provide efficient energy conversion and reduce the velocity substantially below MACH 1 to further improve the efficiency of the subsonic diffusion downstream therefrom. A logarithmic spiral is approximated by a circular arc subtended by the channel longitudinal axes to permit recovery of angular momentum while the circular cross section of the channels permits recovery of swirl velocity energy. The required diameter of the outer circumference of the diffuser is reduced by using the shock waves to greatly reduce gas velocity within a short distance, by the curvature of the channels and by the angle of incidence of the longitudinal channel axes.

Aircraft air intakes

Abstract: A supersonic two-dimensional intake arrangement for two or more jet propulsion engines has a duct for each engine, a mouth region for each duct, said mouth regions being positioned in adjacent rearewardly staggered rearwardly such that at least one shock wave is common to all mouth regions. Preferably one duct wall is common to adjacent intake mouths.

On the dynamics of hypervelocity liquid jet impact on a flat rigid surface

Abstract: The phenomena associated with the impact of a right circular fluid cylinder with a flat rigid surface are studied with the aid of a two-dimensional axisymmetric finite difference code. Both sub-and supersonic initial conditions are investigated. It is shown that, contrary to earlier reports, the maximum pressure sustained is precisely the one-dimensional maximum, i.e., the shock Hugoniot. The relaxation time corresponds to the time for the edge rarefaction, initiating at the impact corner, to traverse the jet radius. The maximum lateral speed, at the impact corner, was found to be nearly twice that of impact, slightly higher for very low Mach number and slightly lower for supersonic impact.

On the application of linear stability theory to the problem of supersonic boundary-layer transition

Abstract: Linear stability theory is used to calculate the amplitude ratio of constant-frequency disturbances as a function of Reynolds number for insulated and cooled-wall flat-plate boundary layers between Mach numbers 1.3 and 5.8. The growth curves are used to examine the consequences of using a fixed amplitude of the most unstable frequency as a transition criterion. The effect of free-stream Mach number on insulated-wall boundary layers is calculated, assuming that the initial disturbance level is constant, is proportional to the square of the free-stream Mach number and to the square root of the energy density of the one-dimensional power spectra of free-stream disturbances measured in supersonic wind tunnels.

Recent Contributions to Experiments on Cylindrical Shell Panel Flutter

Abstract: Experimental studies have been carried out on the aeroelastic stability of several cylindrical shell configurations under various internal stress levels and supersonic flow conditions. These experiments were conducted in the AEDC Propulsion Wind Tunnel Facility over the Mach number range 1.2-3.5. They demonstrated that the still-air buckling characteristics of thin cylindrical shells were not significantly influenced by the supersonic airstream. In addition, two basic types of panel flutter instabilities were found. One was of a mild limited amplitude motion which resulted in no apparent damage to the shell even though it persisted for several minutes. The second type of panel flutter encountered was more highly divergent or explosive in character destroying the shell within a few seconds after its onset. The nature of these two types of panel flutter instabilities were found to be closely linked to the fluid boundary-layer characteristics.

Device for measuring three dimensional coordinates of models

Abstract: A device for measuring three dimensional coordinates of models including a supersonic oscillator T placed at a point P on models to be measured, at least three supersonic receivers A, B and C provided at three known points in the coordinates and means for measuring, electrically, the time required for the supersonic waves to travel from the supersonic oscillating point t of said oscillator T to the supersonic receiving points a, b and c of the receivers, A, B and C, respectively. This period of time is converted to distance, and the three dimensional coordinates of the points to be measured are computed, based on the coordinates of the said known points and the converted distance.

On the Unsteady Supersonic Cascade With a Subsonic Leading Edge—An Exact First Order Theory—Part 1

Abstract: This paper presents, in two parts, the theoretical predictions of the aerodynamic forces acting on slowly oscillating airfoils in a supersonic cascade with a subsonic leading edge. The analysis is based on the assumption of an inviscid, two-dimensional and linearized flow. In the first part of the paper, the flow field ahead of the cascade is considered. An initial value problem is posed and, from the periodicity requirement in the cascade, the problem is reformulated in terms of integral equations. Solution of the integral equations, accurate to the first order of a frequency parameter, are obtained in closed form. In the limit of the steady flow, the unsteady flow analysis yields a mathematical verification of the unique incidence effect. Based on this proof, a simple rule is presented for the airfoil suction surface contour satisfying steady flow requirement ahead of the cascade. The complete aeroelastic problem, including the solution for the flow field between the blades and the trailing interference zone, is treated in Part 2.