Showing papers on "Supersonic speed published in 1975"

High-Lift Aerodynamics

Abstract: c. f = chord fraction, see Eq. (5.1) H = shape factor of the boundary layer, d*/0 £ = plate length L = lift m = exponent in Cp=x flows, also lift magnification factor (5.1) M = Mach number p = pressure q = dynamic pressure Q = flow rate R = Reynolds number (= u Ox/v in Stratford flows) R6 = Reynolds number based on momentum thickness uee/v S = Stratford's separation constant (4.10); also peripheral distance around a body or wing area / = blowing slot gap, also thickness ratio of a body u = velocity in x-direction u0 = initial velocity at start of deceleration in canonical and Stratford flows v = velocity normal to the wall V = a general velocity x = length in flow direction, or around surface of a body measured from stagnation point if used in connection with boundary-layer flow

Wind Tunnel Experiments Relating to Supersonic and Hypersonic Boundary-Layer Transition

Abstract: Hot-wire anemometry is used to study the origin and growth of "natural" fluctuations in zero pressure-gradient boundary layers for several Mach numbers between 1.6 and 8.5. The importance to transition of certain physical mechanisms is examined through comparison of the fluctuation growth with the sound-forcing and stability theories of Mack. Flow fluctuations of substantial amplitude were observed within the laminar layer ahead of stations where instability amplification is expected to be important. These fluctuations were found to be cross-correlated with the sound field for the higher supersonic speeds, but not for the lower ones. The fluctuation growth rates in the unstable Reynolds number range ahead of the nonlinearity region were in reasonably close agreement with the theory for Mach 4.5; the agreement for Mach 2.2 and 8.5 was qualitative. The second mode of instability was predominant at Mach 8.5.

Steady and Oscillatory Subsonic and Supersonic Aerodynamics around Complex Configurations

Abstract: A general formulation for steady and oscillatory, subsonic and supersonic, potential linearized aerodynamic flows around complex configurations is presented. A linear integral equation relating the unknown potential on the surface of the body to the known downwash is used. The formulation is applied to the analysis of flowfields around wings and wing-body combinations. The surface is divided into small quadrilateral elements which are approximated with hyperboloidal surfaces. The potential is assumed to be constant within each element. This yields a set of linear algebraic equations. The coefficients are evaluated analytically. Numerical results for steady and oscillatory, subsonic and supersonic flows indicate that the method, is not only more general and flexible than other available methods, but is also fast, accurate, and in excellent agreement with existing results.

Experiments on the instability waves in a supersonic jet and their acoustic radiation

Abstract: An experimental investigation of the instability and the acoustic radiation of the low Reynolds number axisymmetric supersonic jet has been performed. Hot-wire measurements in the flow field and microphone measurements in the acoustic field were obtained from different size jets at Mach numbers of about 2. The Reynolds number ranged from 8000 to 107000, which contrasts with a Reynolds number of 1·3 × 106 for similar jets exhausting into atmospheric pressure.Hot-wire measurements indicate that the instability process in the perfectly expanded jet consists of numerous discrete frequency modes around a Strouhal number of 0·18. The waves grow almost exponentially and propagate downstream at a supersonic velocity with respect to the surrounding air. Measurements of the wavelength and wave speed of the St = 0·18 oscillation agree closely with Tam's theoretical predictions.Microphone measurements have shown that the wavelength, wave orientation and frequency of the acoustic radiation generated by the dominant instability agree with the Mach wave concept. The sound pressure levels measured in the low Reynolds number jet extrapolate to values approaching the noise levels measured by other experimenters in high Reynolds number jets. These measurements provide more evidence that the dominant noise generation mechanism in high Reynolds number jets is the large-scale instability.

Advances in fan and compressor blade flutter analysis and predictions

Abstract: A unified approach to flutter prediction has been developed at Pratt & Whitney Aircraft (P&WA). The aeromechanical stability of the blade-disk system is expressed in terms of a stability parameter which measures the amount of unsteady work done by the air on the system, when the system is vibrating in one of its natural modes. In neutrally stable systems, the unsteady work done by the air on the blades will balance the work dissipated by friction and by material damping. An accurate prediction of the vibrational deflections and of the unsteady aerodynamic forces is required at every spanwise location on each blade, so that the work done by the unsteady aerodynamic forces may be calculated. Recent progress is described in the prediction of unsteady aerodynamic forces and the determination of mode shapes. The stability model is applied to the prediction of supersonic flutter, chord wise bending flutter, and stall flutter. Recommendations are made for additional work necessary to improve the prediction model.

Two-Dimensional Shock Wave-Boundary Layer Interactions in High Speed Flows,

Abstract: : A review is made of two-dimensional supersonic interactions including separation for laminar and turbulent flows. Part 1 discusses recent theoretical developments in interacting flows and presents numerical techniques for calculating these flows, including finite difference and integral methods. Theoretical discussions are presented for both laminar and turbulent interactions. Part 2 reviews recent experimental studies which have been directed towards understanding the fluid mechanics of attached and separated regions of shock wave-boundary layer interaction in the supersonic annd hypersonic flow.

Control of shock-wave boundary-layer interactions by bleed in supersonic mixed compression inlets

Abstract: An experimental investigation has been conducted to determine the effect of bleed region geometry and bleed rate on shock wave-boundary layer interactions in an axisymmetric, mixed-compression inlet at a Mach number of 2.5. The full realizable reduction in transformed form factor is obtained by bleeding off about half the incident boundary layer mass flow. Bleeding upstream or downstream of the shock-induced pressure rise is preferable to bleeding across the shock-induced pressure rise. Slanted holes are more effective than normal holes. Two different bleed hole sizes were tested without detectable difference in performance.

Calculation of Muzzle Blast Flowfields

Abstract: The muzzle blast field generated by a gun-launched high-velocity projectile is characterized by a highly underexpanded supersonic exhaust plume, which terminates at a strong shock (the Mach disk), an expanding front of exhaust gases (the contact surface), and an expanding, nearly spherical outer shock (the blast wave). The present study is directed toward theoretical description of the inviscid gas dynamics of the blast field. The rioted features are discussed in terms of well-establis hed theories for spherical blast waves with variable energy release and for steady underexpanded plumes, from which their interaction can be qualitatively described. To obtain a quantitative representation, a finite difference solution is developed for the unsteady compressible flow between the Mach disk and the blast wave, assuming spherical symmetry. The results obtained are in good agreement with experimental measurements of the motion of the blast wave, the contact surface and Mach disk for a 3200 fps round fired from an M16 rifle. a B C d E

High energy laser beam cutting method and apparatus

Abstract: Method and device enabling the injecting in the cutting groove of a gas coming from a nozzle at a supersonic speed, brought to a temperature which may probably be very high. A device for injecting an auxiliary gas ensures the protection of the front lens of the laser. Increase in the cutting speed.

Computation of three-dimensional, inviscid supersonic flows

Abstract: The paper sets forth in detail a method for the finite-difference computation of three-dimensional supersonic fields in an Eulerian mesh. First-, second-, and third-order finite difference schemes are examined. Attention is given to proper treatment of the impermeable and permeable boundaries encompassing the computational plane. Numerical results are presented for certain specific configurations: a conical wing-body combination, internal corner flow, a two-dimensional blunt body, an interfering shock problem, and three-dimensional inviscid supersonic flow past a shuttle-orbiter type vehicle.

Boundary-Layer Transition on Supersonic Cones in an Aeroballistic Range

Abstract: Research was undertaken with the purpose of determining the effect of the unit Reynolds number on boundary-layer transition under conditions where disturbances associated with wind tunnel flows would not be present. The location of boundary-layer transition was determined from shadowgrams of nominally sharp, 4° and 10° semi-angle cones in an aeroballistic range at freestream Mach numbers of 23 and 5.0 and unit Reynolds numbers of 0.3 x 10 to 8 x 10 per in. Owing to constant and equal freestream and cone skirt temperatures, the average ratio of cone wall-to-adiabatic recovery temperature was 0.52 at Mach 2.3 and 0.19 at Mach 5.0. Features of free-flight experimentation that may be suspected of influencing boundary-layer transition were investigated. These included 1) oscillatory motion and finite angles of attack, 2) surface roughness, 3) vibration of the model, and 4) non-uniform (hot-tip) surface temperature. There was no evidence that any of these conditions influenced the major results. The data show local Reynolds number of transition increasing with unit Reynolds number for both Mach numbers. A siren was used to elevate the fluctuating sound pressure ratio by a factor of 200, but that produced no measurable effect on transition locations.

Flutter of Flat Finite Element Panels in a Supersonic Potential Flow

Abstract: A finite element formulation and a solution procedure are developed for the flutter analysis of twodimensional flat panels with one surface exposed to a supersonic potential flow and supported at the leading and trailing edges. The aerodynamic forces due to supersonic flow are obtained from the exact theory for linearized two-dimensional unsteady flow where no limitations on the order of the frequency are imposed. Thus the Mach number need not be limited to beyond approximately 1.6. When using this theory, the aerodynamic matrix becomes location-dependent and must be formulated for each element. This difficulty is overcome by first assembling the element shape functions for the entire structure and then finding the aerodynamic forces and the element aerodynamic matrix by numerical integration. The effects of structural damping and initial inplane tension are included. Examples for finding the panel thickness required to prevent flutter at various Mach numbers are demonstrated and the results are compared with an alternative Galerkin's modal solution as well as experiments.

Unsteady Supersonic Cascade in Subsonic Axial Flow

Abstract: This paper presents an analysis for determining the unsteady flowfield produced by an oscillating cascade placed in a supersonic stream which has a subsonic velocity component normal to the cascade. The analysis is based on the assumptions of an inviscid, two dimensional, linearized flowfield. Solutions for the velocity potential and the blade pressure distributions which satisfy the blade-to-blade periodicity condition are developed explicitly in terms of disturbance functions distributed on blade and wake surfaces. The boundary conditions of flow tangency at blade surfaces and continuity of pressure across wake surfaces provide integral relations which can be solved numerically to evaluate the disturbance functions. Predicted blade pressure distributions are in good agreement with results determined from a previous finite cascade solution. Further, in the two limiting cases of sonic axial velocity and zero frequency, the present solution approaches the lower limit of Lane's solution for supersonic axial flow, and it reduces to an Ackeret type of steady-state solution, respectively. The numerical examples indicate that a single~degree- of-freedom torsional instability will exist over a broad range of cascade parameter values.

Supersonic wave flaw detecting apparatus

Abstract: A supersonic or ultrasonic wave flaw detector in which acoustic boundaries of flaws and/or abnormality of materials within an object to be inspected is detected by supersonic or ultrasonic waves and the position and shape of such a flaw and/or abnormality are displayed. A hard object to be inspected is sector-scanned with a supersonic wave beam with a variable-angle beam probe in direct contact with the object. In a range of comparatively small incident angles at which the longitudinal wave mainly penetrates the object, sectional images due to echo signals of the longitudinal wave are displayed on a display screen with scanning lines representative of the transmission speed and direction of the longitudinal wave. In the range of large incident angles where the shear wave penetrates the object, on the other hand, sectional images due to the echo signals of the shear wave are displayed on the display screen with scanning lines representative of the transmission speed and direction of the shear wave. For this, the sweeping speed and direction for the display scanning lines are switched in accordance with the ranges of incident angles. The switched flaw detecting operation modes of longitudinal and shear waves thus combined covers a complete sectional image to be displayed on the display screen.

Experiments on supersonic jet noise

Abstract: The noise generated by a 1-in. supersonic jet was investigated at nominal jet Mach numbers of 1.5, 2.0, and 2.5. In particular, a quantity W, referred to as the apparent source strength per unit length, was determined along the jet axis using a directional microphone system. The integrated value of W along the jet axis was found to agree with the sound intensity obtained by a conventional microphone. This result is consistent with the a priori assumption that the jet may be described in terms of independent, spatially compact acoustic sources. The main finding of the investigation is the discovery of two distinct intense noise-producing regions in a jet having supersonic source velocities: the upstream region associated with Mach wave radiation, and a zone, located downstream of the potential cone, exhibiting radiation similar in character to that of a subsonic jet. An estimate of the radiated intensity associated with the Mach waves also is made.

Design and performance at a local Mach number of 6 of an inlet for an integrated Scramjet concept

Abstract: The development of a concept for a modular supersonic combustion ramjet which is designed to integrate with the airframe of a hypersonic vehicle is presented. The design philosophy and results of experiments at Mach 6 to evaluate the performance of the scramjet inlet are given. The inlet was designed with modest contraction ratio, fixed geometry, and three fuel injection struts which contributed to the inlet flow compression and provided a short combustor design that resulted in low internal cooling requirements. Results indicate that the inlet performance is well within the acceptable range for high engine performance.

Minimization of sonic-boom parameters in real and isothermal atmospheres

Abstract: The procedure for sonic-boom minimization introduced by Seebass and George for an isothermal atmosphere was converted for use in the real atmosphere by means of the appropriate equations for sonic-boom pressure signature advance, ray-tube area, and acoustic impedance. Results of calculations using both atmospheres indicate that except for low Mach numbers or high altitudes, the isothermal atmosphere with a scale height of 7620 m (25 000 ft) gives a reasonable estimate of the values of overpressure, impulse, and characteristic overpressure obtained by using the real atmosphere. The results also show that for aircraft design studies, propagation of a known F-function, or minimization studies at low supersonic Mach numbers, the isothermal approximation is not adequate.

Two-dimensional jet interaction studies at large values of Reynolds and Mach numbers

Abstract: The situation chosen for study is a gaseous jet that is injected through a transverse slot nozzle in a wall and into a supersonic external flow which is uniform outside of a turbulent boundary layer. Experiments were conducted with normal, sonic jets and forward-facing steps at external flow Mach numbers of 2.5 to 13, and Reynolds numbers based on running length of 7.5 X 10 to 5.5 x 10. The amplification factor (the upstream interaction force plus the jet thrust normalized by the vacuum thrust of a sonic jet) is relatively insensitive to variations in external flow Mach number and Reynolds number. The effect of pressure ratio on amplification factor is very small when the external flow properties and jet mass flow rate are held constant. Plateau pressures associated with separation upstream of the jet or step, and wall static pressure distributions near the separation line are in good agreement with existing correlations.

Numerical solutions of supersonic and hypersonic laminar flows over a two-dimensional compression corner

Abstract: An efficient time-splitting, second-order accurate, numerical scheme is used to solve the complete Navier-Stokes equations for supersonic and hypersonic laminar flow over a two-dimensional compression corner. A fine, exponentially stretched mesh spacing is used in the region near the wall for resolving the viscous layer. Good agreement is obtained between the present computed results and experimental measurement for a Mach number of 14.1, a Reynolds number of 104,000, and wedge angles of 15, 18, and 24 deg. The details of the pressure variation across the boundary layer are given, and a correlation between the leading edge shock and the peaks in surface pressure and heat transfer is observed.

Small Disturbance Transonic Flows About Oscillating Airfoils and Planar Wings.

Abstract: : A theory and numerical solution method are presented for the problem of two or three dimensional flow about thin wings undergoing harmonic oscillations. The theory is based on a treatment of the unsteady flow as a small perturbation on the nonlinear steady flow. The coupled governing equations for the steady and unsteady perturbation potentials are of mixed elliptic/hyperbolic type and are solved using the mixed differencing line relaxation techniques of Murman and Cole. The theory and solution method have been generalized to treat subsonic or supersonic freestream flows, solid or porous wall effects and three dimensional flow about planar wings. Calculations demonstrating each of these capabilities are presented. Accuracy of the method for realistic steady and unsteady flows is examined by comparison to more exact numerical results and experimental data. The comparison indicates that the present method provides a useful and efficient inviscid solution. Quantitative differences between the present results and experimental data are attributed primarily to viscous effects, pointing out the importance of such effects for transonic flows.

Generalized unsteady embedded Newtonian flow

Abstract: An analysis is presented which extends the previously developed unsteady embedded Newtonian theory down to finite supersonic Mach numbers. It is found that Mach number can have a very large influence on the stability characteristics of slender blunted cones, and that there exist cone angle-nose bluntness combinations for which these Mach number effects are minimized. The computed effects of nose bluntness on static and dynamic stability derivatives are in excellent agreement with available experimental data. This also holds true for the highly nonlinear effects of angle of attack.

Effect of Injector Shape on Penetration and Spread of Liquid Jets

Abstract: L injection into a supersonic air stream finds applications in supersonic combustion ramjets (scramjets), transpiration cooling of re-entry bodies and thrust vector control of rockets. The injector geometry will have a significant role in liquid injection applications. The injection of liquid into a supersonic air stream produces an interaction shock and a freestream boundary-layer separation zone upstream of the injector. The separation zone plays an important role during combustion due to the high rate of heat transfer to the wall in this region. The shock system associated with each injector in a practical supersonic combustor has two important effects: 1) it reduces the total pressure of the freestream and thus adversely affects the overall performance of the engine; 2) static temperature and pressure of the freestream rise through the injector shock system thus creating better conditions from the viewpoint of chemical reaction rates. The shape and strength of the shock also affect the forces on the liquid column and thus penetration. In general, the shock system is a strong function of injector geometry. The influence of injector geometry on penetration and spread of the jet were given major emphasis in the present investigation. The purpose was to obtain experimental penetration and spread data suitable for engineering use and to seek theoretical correlations incorporating and governing parameters. The motivation for the present work comes from the work of Kush and Schetz who observed that a liquid jet through a rectangular slot aligned with the flow gives significantly higher penetration than through a circular hole of the same area.

On the flow near a weak shock wave downstream of a nozzle throat

Abstract: : In a transonic nozzle flow for which the velocity is slightly supersonic in some neighborhood of the nozzle throat, a shock wave may be present either very close to the throat or else somewhat further downstream. In the latter case, relatively simple series solutions in general provide an asymptotic description of the fluid motion except very close to the shock wave. These outer solutions are reviewed for symmetric two-dimensional flow, and a correction is derived in the form of an inner solution for a small region immediately behind the shock. The resulting solution exhibits the characteristics which are expected at the intersection of a normal shock wave and a curved wall. An extension to axisymmetric flow is also given. (Modified author abstract)

The electron-beam fluorescence technique for measurements in hypersonic turbulent flows

Abstract: The electron-beam fluorescence technique has been employed in measuring simultaneous density and temperature fluctuations in a hypersonic (M ≃ 16), adiabatic wall boundary layer. The paper discusses this technique, as it is applied to the conditions of relatively high density associated with turbulent flows. It presents general considerations concerning the attainable frequency response and spatial resolution of the technique. It describes results from initial measurements in a boundary layer on a wind-tunnel wall. These results show that the r.m.s. of the density, temperature and pressure fluctuations are large, much larger than observed in supersonic boundary layers.