Showing papers on "Supersonic speed published in 1988"

Dynamic pressure loads associated with twin supersonic plume resonance

Abstract: The phenomenon of twin supersonic plume resonance is defined and studied as it pertains to high level dynamic loads in the inter-nozzle region of aircraft like the F-15 and B1-A. Using a 1/40th scale model twin jet nacelle with powered choked nozzles, it is found that intense internozzle dynamic pressures are associated with the synchrophased coupling of each plume's jet flapping mode. This condition is found most prevalent when each plume's jet flapping mode has constituent elements composed of the B-type helical instability. Suppression of these fatigue bearing loads was accomplished by simple geometric modifications to only one plume's nozzle. These modifications disrupt the natural selection of the B-type mode and thereby decouple the plumes.

Airfoil dynamic stall at constant pitch rate and high Reynolds number

Abstract: An experiment has been performed to study the aerodynamics of dynamic stall penetration at constant pitch rate and high Reynolds number, in an attempt to model more accurately conditions during aircraft poststall maneuvers and during helicopter high-speed forward flight. An airfoil was oscillated at pitch rates, A = ac/2U between 0.001 and 0.020, Mach numbers between 0.2 and 0.4, and Reynolds numbers between 2-4 x 10. Surface pressures were measured using 72 miniature transducers, and the locations of transition and separation were determined using 8 surface hot-film gages. The results demonstrate the influence of the leading-edge vorticity on the unsteady aerodynamic response during and after stall. The vortex is strengthened by increasing the pitch rate and is weakened by increasing the Mach number and by starting the motion close to the steady-state stall angle. A periodic pressure oscillation occurred after stall at high pitch angle and moderate Reynolds number; the oscillation frequency was close to that predicted for a von Karman vortex street. A small supersonic zone near the leading edge at M = 0.4 was found to reduce significantly the peak suction pressures and the unsteady increments to the airloads. These results provide the first known data base of constant-pitch-rate aerodynamic information at realistic combinations of Reynolds and Mach numbers.

Observations of supersonic free shear layers

Abstract: Visual spreading rates of turbulent shear layers with at least one stream supersonic were measured using Schlieren photography. The experiments were done at a variety of Mach number-gas combinations. The spreading rates are correlated with a compressibility-effect parameter called the convective Mach number. It is found that for supersonic values of the convective Mach number, the spreading rate is about one quarter that of an incompressible layer at the same velocity and density ratio. The results are compared with other experimental and theoretical results.

Efficient algorithm for solution of the unsteady transonic small-disturbance equation

Abstract: A time accurate approximation factorization (AF) algorithm is formulated for solution of the three-dimensional unsteady transonic small-disturbance equation. The AF algorithm consists of a time linearization procedure coupled with a Newton iteration technique. Superior stability characteristics of the new algorithm are demonstrated through applications to steady and oscillatory flows at subsonic and supersonic freestream conditions for an F-5 fighter wing. For steady flow calculations, the size of the time step is cycled to achieve rapid convergence. For unsteady flow calculations, the AF algorithm is sufficiently robust to allow the step size to be selected based on accuracy rather than on stability considerations. Therefore, accurate solutions are obtained in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods.

Combustion Processes in Supersonic Flow

Abstract: Three general classes of models that describe the processes occurring in diabatic flow in ducts having supersonic entry conditions are discussed. They are: integral techniques, finite-difference methods, and exact two-dimensional planar flame models formulated on the basis of instantaneous heat release. All three methods rigorously satisfy the conservation equations. The first two methods provide a basis for predicting and analyzing supersonic combustor performance. The careful interpretation and judicious use of experimental observations are crucial for the successful application of these methods. Comparisons of analytical and experimental results are presented, and generalized parametric studies are included. The third method is based on an idealized mixing and combustion model that may not be achievable, but nonetheless serves as a valuable analytical tool for explaining complex processes involving shock waves and heat addition. Results from four types of flow structures are discussed. Nomenclature A = cross-sectional area Af = projected area of inlet Aw = wall area

A two-dimensional numerical simulation of a supersonic, chemically reacting mixing layer

Abstract: Research has been undertaken to achieve an improved understanding of physical phenomena present when a supersonic flow undergoes chemical reaction. A detailed understanding of supersonic reacting flows is necessary to successfully develop advanced propulsion systems now planned for use late in this century and beyond. In order to explore such flows, a study was begun to create appropriate physical models for describing supersonic combustion, and to develop accurate and efficient numerical techniques for solving the governing equations that result from these models. From this work, two computer programs were written to study reacting flows. Both programs were constructed to consider the multicomponent diffusion and convection of important chemical species, the finite rate reaction of these species, and the resulting interaction of the fluid mechanics and the chemistry. The first program employed a finite difference scheme for integrating the governing equations, whereas the second used a hybrid Chebyshev pseudospectral technique for improved accuracy.

Shock-wave-turbulent-boundary-layer interaction and its control - A survey of recent developments

Abstract: This paper presents an overview of some of the recent developments that have taken place in the understanding, prediction and control of two-dimensional shock-wave-turbulent-boundary-layer interaction at high speeds. Following a brief description of the upstream influence phenomena, detailed discussions of incipient and fully separated flows at supersonic and transonic speeds are presented. A brief account of certain gross unsteady features of shock-separated flows is given next. Typical examples demonstrating the current ability to predict these complex flows are also included. Finally, a review of techniques using suction and tangential blowing for controlling shock-separated flows is presented.

An asymptotic analysis of supersonic reacting mixing layers

Abstract: The purpose of this paper is to present an asymptotic analysis of the laminar mixing of the simultaneous chemical reaction between parallel supersonic streams of two reacting species. The study is based on a one-step irreversible Arrhenius reaction and on large activation energy asymptotics. Essentially it extends the work of Linan and Crespo to include the effect of free shear and Mach number on the ignition regime, the deflagration regime and the diffusion flame regime. It is found that the effective parameter is the product of the characteristic Mach number and a shear parameter.

Aircraft engine inlet cowl anti-icing system

Abstract: A system is provided for preventing and removing ice accumulation from the leading edges of an inlet cowl of an aircraft gas turbine engine. High pressure, high temperature air, preferrably from a port on the engine compressor, is provided to an annular duct located within an annular chamber formed at the leading edge of the inlet cowl. A plurality of air ejector nozzles are spaced about the air duct and function to direct anti-icing air at a supersonic velocity toward the forward inboard surface of the inlet cowl and create a swirling air mass within the annular chamber.

Computational analysis of unsteady supersonic cavity flows driven by thick shear layers

Abstract: Supersonic cavity flows driven by a thick shear layer at Mach 1·5 and 2·5 are studied by solving the two-dimensional unsteady compressible Navier-Stokes equations in terms of mass-averaged variables. The length to depth ratio of the rectangular cavity is three. The numerical scheme used is the finite-difference algorithm by Brailovskaya. A two-layer eddy-viscosity turbulence model is used. The results are compared with experimental data. The computations show the self-sustained oscillations at Mach 1·5 and 2·5. The continuous formation and downstream shedding of leading edge vortices is demonstrated. The oscillatory modes are correctly predicted. The first mode is attributed to a large unsteady trailing edge vortex moving in the transverse direction. Based on the analysis, it is considered that the oscillation in the length to depth ratio three cavity is a longitudinal one and is controlled by a fluid dynamic mechanism rather than a purely acoustic one.

Passive venting system for modifying cavity flowfields at supersonic speeds

Abstract: The drag of airfoils in transonic flow can be reduced through the use of a passive venting system that employs a porous plate for part of the airfoil upper surface with a vent chamber underneath the porous plate Attention is given to the results obtained with a wind tunnel model employing such a porous floor system. This passive venting system has been used to extend the length/height value before the onset of high drag-producing closed cavity flow at supersonic speeds.

2-D transonic flow with energy supply by homogeneous condensation: onset condition and 2-D structure of steady Laval nozzle flow

Abstract: A generalized form of the similarity law for the condensation onset Mach number of water vapor in air in the transonic and supersonic range for water vapor flow in moist air is derived from well known basic approaches for supersonic nozzles. These statements are confirmed by extensive experimental investigations in Laval nozzles, as well as by results of other authors and computations on the basis of the Euler equation linked with the classical theory of nucleation and droplet growth. In this experimental research priority is given to the qualitative description of the two-dimensional condensation processes, and their effects in transonic flows in nozzles of different geometrical configuration (e. g. slightly or well curved). A quantitative discussion of 2-D structures in condensation regions requires the introduction of a characteristic angle along streamlines. It is then directly possible to describe the different types of compression disturbances in supersonic flows with heat addition.

Experimental Investigation of the Performance of a Supersonic Compressor Cascade

Abstract: Results are presented from an experimental investigation of a linear, supersonic, compressor cascade tested in the supersonic cascade wind tunnel facility at the DFVLR in Cologne, Federal Republic of Germany. The cascade design was derived from the near-tip section of a high-through-flow axial flow compressor rotor with a design relative inlet Mach number of 1.61. Test data were obtained over a range of inlet Mach numbers from 1.23 to 1.71, and a range of static pressure ratios and axial-velocity-density ratios (AVDR) at the design inlet condition. Flow velocity measurements showing the wave pattern in the cascade entrance region were obtained using a laser transit anemometer. From these measurements, some unique-incidence conditions were determined, thus relating the supersonic inlet Mach number to the inlet flow direction. The influence of static pressure ratio and AVDR on the blade passage flow and the blade-element performance is described, and an empirical correlation is used to show the influence of these two (independent) parameters on the exit flow angle and total-pressure loss for the design inlet condition.

Unique, clean-air, continuous-flow, high-stagnation-temperature facility for supersonic combustion research

Abstract: Accurate, spatially-resolved measurements can be conducted of a model supersonic combustor in a clean air/continuous flow supersonic combustion facility whose long run times will allow not only the point-by-point mapping of flow field variables with laser diagnostics but facilitate the simulation of steady-state combustor conditions. The facility will provide a Mach 2 freestream with static pressures in the 1 to 1/6 atm range, and stagnation temperatures of up to 2000 K.

Combined multi-speed jet engine for the drive of airplanes and space vehicles

Abstract: Composite changeover-type reaction power plant for propelling aircraft and spacecraft operating in subsonsic, supersonic and hypersonic regions, embracing a turbofan bypass engine with front fan, and a ramjet engine with a cycle corresponding to that of the fan, where the fan has individually independently supported, operationally counterrotating statorless rotors and variable fan blades which are feathered and immobilized during ramjet operation for low drag during said ramjet operation.

Disruption of galactic radio jets by shocks in the ambient medium

Abstract: Observations show that jets in moderate luminosity (<1025 W Hz−1 at 20 cm wavelength) radio galaxies can flare dramatically in a few jet diameters and with opening angles up to 90° into diffuse lobes or tails1–4 (Fig. 1). These morphologies are difficult to reproduce in numerical simulations of supersonic jets that move outward in constant5,6 or smoothly varying atmospheres (N. Norman, unpublished results). By analogy with structures seen in the laboratory7, one could interpret the collimated jets as moderately supersonic (Mach number 2–5) fluid flows, and the lobes or tails as subsonic plumes that are subject to turbulent broadening and entrainment of the ambient medium. Our studies indicate that the transition from supersonic to subsonic flow can occur suddenly only if a strong planar (normal to the flow direction) shock or Mach disk forms within the jet. Here we show that such an internal shock is produced as the jet crosses a shock wave in the external medium. The external shock could form, for example, by a galactic wind encountering the intergalactic medium. We find that for jet disruption the jet Mach number must be less than the wind-shock Mach number, a result also understandable from analytic arguments. We apply this model to Cen A and wide-angle-tailed radio galaxies.

Combustion in supersonic flow

Abstract: Examples of numeric modeling of complex combustion processes in supersonic flows are presented which illustrate the possibility of obtaining useful information which may reveal their basic regularities determined by the parameters of the entrance streams, mode of injection, type of fuel, and the interaction of various factors pertaining to combustion. Hydrogen combustion was used in the analysis. Peculiar results are presented on numerical modeling in situations where the kinetics of the exothermal reactions were subject to limitations imposed by the hydrodynamic conditions, and where the development of the motion of a continuous medium was part of the combustion process. The discovery of small supersonic velocities in the reverse flow in the recirculation zone, confirmed by the presence of static pressure distributions, was noted and studied.

Unsteady viscous calculations of supersonic flows past deep and shallow three-dimensional cavities

Abstract: Computational simulations were performed for supersonic, turbulent flows over deep and shallow three-dimensional cavities. The width and the depth of these cavities were fixed at 2.5 in. and 0.5 in., respectively. Length-to-depth ratio of the deep cavity was 6 and that of the shallow cavity was 16. Freestream values of Mach number and Reynolds number were 1.50 and 2.0 x 10 to the 6th/ft., respectively, at a total temperature of 585 R. The thickness of the turbulent boundary layer at the front lip of the cavity was 0.2 in. Simulations of these oscillatory flows were generated through time-accurate solutions of Reynolds-averaged full Navier-Stokes equations using the explicit MacCormack scheme. The solutions are validated through comparisons with experimental data. The features of open and closed cavity flows and effects of the third dimension are illustrated through computational graphics.

Thickness noise of a propeller and its relation to blade sweep

Abstract: Linearized acoustic theory is applied to the calculation of the thickness noise produced by a supersonic propeller with sharp leading and trailing edges. The theoretical development is summarized and numerical calculations of the pressure-time waveform are presented. The erratic behaviour of previous time-domain calculations has been completely eliminated by careful numerical treatment of singular points, multiple singular points and nearly singular points that appear in the analysis. This allows a close inspection of the details of the calculated waveform and leads to the discovery of abrupt changes of slope in the pressure-time waveform, produced by singular points entering or leaving the blade at the tip. The behaviour of the pressure-time waveform is shown to be closely related to changes in the retarded rotor shape. Logarithmic singularities in the waveform are shown to be produced by regions on the blade edges that move towards the observer at sonic speed while at the same time having the edge normal to the line joining the source point and the observer. The logarithmic singularities are closely related to the shock waves produced by a swept airfoil in supersonic rectilinear motion, and they can be eliminated throughout the entire flow field by sweeping the rotor so that the Mach-number component normal to the leading and trailing edges is subsonic for all points on the rotor edges.

Design and fabrication requirements for low-noise supersonic/hypersonic wind tunnels

Abstract: A schematic diagram of the new proposed Supersonic Low Disturbance Tunnel (SLDT) is shown. Large width two dimensional rapid expansion nozzles guarantee wide quiet test cores that are well suited for testing models at large angle of attack and for swept wings. Hence, this type of nozzle will be operated first in the new proposed large scale SLDT. Test results indicate that the surface finish of pilot nozzles is critical. The local roughness Reynolds number criteria R sub k is approx. = 10 will be used to specify allowable roughness on new pilot nozzles and the new proposed tunnel. Experimental data and calculations for M = 3.0, 3.5, and 5.0 nozzles give N-factors from 6 to 10 for transition caused by Goertler vortices. The use of N is approx. = 9.0 for the Goertler instability predicts quiet test cores in the new M = 3.5 and M = 6.0 axisymmetric long pilot nozzles that are 3 to 4 times longer than observed in the test nozzles to date. The new nozzles utilize a region of radial flow which moves the inflection point far downstream and delays the onset and amplification of the Goertler vortices.

Noise and Vibration Related to the Patterns of Supersonic Annular Flow in a Pressure Reducing Gas Valve

Abstract: Cause of intense aerodynamic noise and vibration from a contoured type valve is revealed in close relation to the supersonic flow patterns. Simple conical plugs are used in the experiments, and the valve pressure ratio is up to twenty. Four typical patterns of the flow are observed by schlieren photography. In one of these patterns, the jet flow along the plug separates from the wall to form an annular jet impinging on the inner wall of the valve chest. Such flow oscillates significantly in resonance with the acoustic modes of the chest cavity. The radiated noise and the dynamic force acting on the valve stem reach intense levels, dominated by some discrete components of the corresponding frequencies. The mechanisms to generate or to suppress the flow oscillation, leading to the intense noise and vibration, are discussed.

Method and apparatus for launching a projectile at hypersonic velocity

Abstract: A projectile is accelerated to hypersonic velocity in an initially closed barrel of a diameter considerably larger than the projectile diameter which is filled with a compressed fuel-oxidizer mixture. The projectile comprises a conical nose portion, an intermediate portion formed to generate oblique detonation waves, and a tapering tail portion provided with several radial vanes. The projectile is propelled by an initiator gun at supersonic speed through one of the intitially closed ends in the barrel, were the detonation waves cause detonation and combustion of the fuel-oxidizer mixture. The detonation results in a high pressure increase to the rear of the projectile accelerating it along the barrel and shooting it at the reached hypersonic speed through the other, initially closed end of the barrel into the open.

Unsteady Transonic Aerodynamics and Aeroelastic Calculations at Low-Supersonic Freestreams

Abstract: A computational procedure is presented to simulate transonic unsteady flows and corresponding aeroelasticity of wings at low-supersonic freestreams. The flow is modeled by using the transonic small-perturbation theory. The structural equations of motions are modeled using modal equations of motion directly coupled with aerodynamics. Supersonic freestreams are simulated by properly accounting for the boundary conditions based on pressure waves along the flow characteristics in streamwise planes. The flow equations are solved using the time-accurate, alternating-direction implicit finite-difference scheme. The coupled aeroelastic equations of motion are solved by an integration procedure based on the time-accurate, linear-acceleration method. The flow modeling is verified by comparing calculations with experiments for both steady and unsteady flows at supersonic freestreams. The unsteady computations are made for oscillating wings. Comparisons of computed results with experiments show good agreement. Aeroelastic responses are computed for a rectangular wing at Mach numbers ranging from subtransonic to upper-transonic (supersonic) freestreams. The extension of the transonic dip into the upper transonic regime is illustrated.