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At least 955 records · Page 53
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Crossing shock wave-turbulent boundary layer interactions

Three-dimensional interactions between crossing shock waves generated by symmetric sharp fins and a turbulent boundary layer on a flat plate are investigated experimentally and theoretically at Mach number 2.95 and freestream unit Reynolds number 1.96 x 10 to the 7th/ft. The incoming boundary layer has a thickness of 4 mm at the location of the fin leading edges. A comparison of experimental and computational results for two sets of fin angles (11 x 11 and 9 x 9 deg) shows general agreement with regard to surface pressure measurements and surface streamline patterns. The principal feature of the streamline structure is a collision of counterrotating vortical structures emanating from near the fin leading edges and meeting at the geometric centerline of the interaction.

Narayanswami, N.↗
NASA NTRS1995

Nonlinear MHD interactions in some solar coronal structures

The magnetohydrodynamic (MHD) interactions of solar coronal fast shock waves of flare and/or nonflare origin with the boundaries of such structures like the streamer and the coronal hole are considered. These boundaries are treated as MHD tangential discontinuities (TD). Different parameters of the observed corona are used. The general case of oblique interaction is studied. It is shown that a fast shock wave is usually refracted as a fast shock wave inside the coronal streamer. But for a special case of the velocity shear across a TD a slow shock wave may be generated. On the contrary the shock wave is refracted inside the coronal hole, indeed as a slow shock wave. The significance of different effects due to the interaction of fast and slow shock waves with the coronal magnetic field is noticed, especially at the time of coronal mass ejection (CME). Thus slow shock waves may help to 'open' the coronal magnetic field, and fast shock waves - to 'close' it. It is also shown, that an oblique fast MHD coronal shock wave may trigger an instability at the boundary of a streamer considered as a TD. It might have a relation with the observed abrupt process of disappearance of the streamer's boundary in the solar corona.

Grib, S. A.↗
NASA NTRS1965

Radiative shock structure.

Shock wave structure in ionized hydrogen, using Eddington approximation to solve Rankine- Hugoniot jump equation

DIFFUSION WAVE↗
NASA NTRS1971

Diffraction of a plane wave by a three-dimensional corner

By the superposition of the conical solution for the diffraction of a plane pulse by a three dimensional corner, the solution for a general incident plane wave is constructed. A numerical program is presented for the computation of the pressure distribution on the surface due to an incident plane wave of any wave form and at any incident angle. Numerical examples are presented to show the pressure signature at several points on the surface due to incident wave with a front shock wave, two shock waves in succession, or a compression wave with same peak pressure. The examples show that when the distance of a point on the surface from the edges or the vertex is comparable to the distance for the front pressure raise to reach the maximum, the peak pressure at that point can be much less than that given by a regular reflection, because the diffracted wave front arrives at that point prior to the arrival of the peak incident wave.

Ting, L.↗
NASA NTRS1991

Calculation of hypersonic shock structure using flux-split algorithms

There exists an altitude regime in the atmosphere that is within the continuum domain, but wherein the conventional Navier-Stokes equations cease to be accurate. The altitude limits for this so called continuum transition regime depend on vehicle size and speed. Within this regime the thickness of the bow shock wave is no longer negligible when compared to the shock stand-off distance and the peak radiation intensity occurs within the shock wave structure itself. For this reason it is no longer valid to treat the shock wave as a discontinuous jump and it becomes necessary to compute through the shock wave itself. To accurately calculate hypersonic flowfields, the governing equations must be capable of yielding realistic profiles of flow variables throughout the structure of a hypersonic shock wave. The conventional form of the Navier-Stokes equations is restricted to flows with only small departures from translational equilibrium; it is for this reason they do not provide the capability to accurately predict hypersonic shock structure. Calculations in the continuum transition regime, therefore, require the use of governing equations other than Navier-Stokes. Several alternatives to Navier-Stokes are discussed; first for the case of a monatomic gas and then for the case of a diatomic gas where rotational energy must be included. Results are presented for normal shock calculations with argon and nitrogen.

Eppard, W. M.↗
NASA NTRS1972

Shock wave-turbulent boundary layer interactions in rectangular channels.

Interaction regions created by the impingement of full span, externally generated, shock waves on a nozzle wall boundary layer were investigated. Incident shock strength was varied to produce unseparated, incipient, and fully separated flow fields. Significant departures from two-dimensionality were observed over the entire range of shock strengths tested and were identified with sidewall and corner boundary layer effects. However, comparisons of present centerline results with published two-dimensional data, obtained under similar test conditions and geometrical constraints, showed excellent agreement (e.g., incipient separation pressure levels, wall pressure distributions, free-interaction, and scale of the interaction region). This raises some question concerning the degree of two-dimensionality achieved in these previous investigations.

Reda, D. C.↗
NASA NTRS1977

Normal shock wave-turbulent boundary layer interactions in transonic flow near separation

The problem of a normal shock wave impinging on a flat plate, turbulent boundary layer is considered for the case where the external flow is transonic. Asymptotic methods are employed. It is shown that there are two outer regions, including the outer part of the boundary layer and the external flow, in which inviscid flow governing equations hold, and two regions near the wall, in which Reynolds and/or viscous stresses need be taken into account. The solutions in the outer regions lead to the pressure distribution on the wall, for which an analytical expression is presented, valid under those conditions when separation is imminent but has not yet occurred. The solutions valid in the inner regions lead to a separation criterion.

Adamson, T. C., Jr.↗
NASA NTRS1996

Experimental Investigation of Crossing Shock Wave-Turbulent Boundary Layer-Bleed Interaction

Results of an experimental investigation of a symmetric crossing shock wave/turbulent boundary layer/bleed interaction are presented for a freestream unit Reynolds number of 1.68 x 10(exp 7)/m, a Mach number of 2.81, and deflection angles of 8 degrees. The data obtained in this study are bleed mass flow rate using a trace gas technique, qualitative information in the form of oil flow visualization, flow field Pitot pressures, and static pressure measurements using pressure sensitive paint. The main objective of this test is two-fold. First, this study is conducted to explore boundary layer control through mass flow removal near a large region of separated flow caused by the interaction of a double fin-induced shock wave and an incoming turbulent boundary layer. Also, a comprehensive data set is needed for computational fluid dynamics code validation.

Kim, Hyun↗
NASA NTRS1985

Instabilities in decelerating supersonic flows with applications to cosmic ray shocks

The nature of instabilities in cosmic ray shocks is investigated by using two distinct models for the shock wave. For wavelengths which are short relative to the thickness of the shock wave, the shock is treated as a smoothly decelerating low, and an appropriate JWKB type expansion is used to describe the perturbations to the flow. In this, the short wavelength regime, the presence of squeezing and an effective g renders strong cosmic ray shocks unstable in a way which is similar to instabilities in other supersonic flows, such as in de Laval nozzle flow or a heat conduction dominated shock wave. In the long wavelength limit, where the shock is treated as a discontinuous transition, a stability function is derived which, if negative, corresponds to unstable disturbances growing exponentially in time. In this case, it was found that if the cosmic ray fluid is relativistic (gamma sub c = 4/3) and the background plasma ideal (gamma = 5/3), then strong shocks are unstable.

Zank, A. P.↗
NASA NTRS1990

Lyman alpha radiation in external galaxies

The Ly alpha line of atomic hydrogen is often a luminous component of the radiation emitted by distant galaxies. Except for those galaxies which have a substantial central source of non-stellar ionizing radiation, most of the Ly alpha radiation emitted by galaxies is generated within regions of the interstellar medium which are photoionized by starlight. Conversely, much of the energy radiated by photoionized regions is carried by the Ly alpha line. Only hot, massive stars are capable of ionizing hydrogen in the interstellar medium which surrounds them, and because such stars are necessarily short-lived, Ly alpha emission traces regions of active star formation. Researchers argue that the strength of the Ly alpha emission observed from external galaxies may be used to estimate quantitatively the dust content of the emitting region, while the Ly alpha line profile is sensitive to the presence of shock waves. Interstellar dust particles and shock waves are intimately associated with the process of star formation in two senses. First, both dust particles and shock waves owe their existence to stellar activity; second, they may both serve as agents which facilitate the formation of stars, shocks by triggering gravitational instabilities in the interstellar gas that they compress, and dust by shielding star-forming molecular clouds from the ionizing and dissociative effects of external UV radiation. By using Ly alpha observations as a probe of the dust content in diffuse gas at high redshift, we might hope to learn about the earliest epochs of star formation.

Neufeld, David A.↗
NASA NTRS1991

Multiple shock-shock interference on a cylindrical leading edge

The details of an experimental study of shock wave interference heating on a cylindrical leading edge representative of the cowl of a rectangular hypersonic engine inlet are presented. This Mach 8 study has provided the first detailed pressure and heat transfer rate distributions on a cylinder resulting from a two-dimensional shockwave interference pattern created by two incident oblique shock waves intersecting the cylinder bow shock wave. The peak heat transfer rate was 38 times the undisturbed flow stagnation point level and occurred when the two oblique shock waves coalesced prior to intersecting the cylinder bow shock wave. Development of pressure deflection diagrams identified a new interference pattern consisting of concomitant supersonic jets separated from each other by a shear layer and submerged in the subsonic region between the bow shock wave and body.

Wieting, Allan R.↗
NASA NTRS1985

Performance predictions for an SSME configuration with an enlarged throat

The Two Dimensional Kinetics (TDK) computer program that was recently developed for NASA was used to predict the performance of a Large Throat Configuration of the Space Shuttle Main Engine (SSME). Calculations indicate that the current design SSME contains a shock wave that is induced by the nozzle wall shape. In the Large Throat design an even stronger shock wave is predicted. Because of the presence of this shock wave, earlier performance predictions that have neglected shock wave effects have been questioned. The JANNAF thrust chamber performance prediction procedures given in a reference were applied. The analysis includes the effects of two dimensional reacting flow with a shock wave. The effects of the boundary layer with a regenatively cooled wall are also included. A Purdue computer program was used to compute axially symmetric supersonic nozzle flows with an induced shock, but is restricted to flows with a constant ratio of specific heats. Thus, the TDK program was also run with ths assumption and the results of the two programs were compared.

Nickerson, G. R.↗
NASA NTRS2004

Exhaust Nozzle for a Multitube Detonative Combustion Engine

An improved type of exhaust nozzle has been invented to help optimize the performances of multitube detonative combustion engines. The invention is applicable to both air-breathing and rocket engines used to propel some aircraft and spacecraft, respectively. In a detonative combustion engine, thrust is generated through the expulsion of combustion products from a detonation process in which combustion takes place in a reaction zone coupled to a shock wave. The combustion releases energy to sustain the shock wave, while the shock wave enhances the combustion in the reaction zone. The coupled shockwave/reaction zone, commonly referred to as a detonation, propagates through the reactants at very high speed . typically of the order of several thousands of feet per second (of the order of 1 km/s). The very high speed of the detonation forces combustion to occur very rapidly, thereby contributing to high thermodynamic efficiency. A detonative combustion engine of the type to which the present invention applies includes multiple parallel cylindrical combustion tubes, each closed at the front end and open at the rear end. Each tube is filled with a fuel/oxidizer mixture, and then a detonation wave is initiated at the closed end. The wave propagates rapidly through the fuel/oxidizer mixture, producing very high pressure due to the rapid combustion. The high pressure acting on the closed end of the tube contributes to forward thrust. When the detonation wave reaches the open end of the tube, it produces a blast wave, behind which the high-pressure combustion products are expelled from the tube. The process of filling each combustion tube with a detonable fuel/oxidizer mixture and then producing a detonation repeated rapidly to obtain repeated pulses of thrust. Moreover, the multiple combustion tubes are filled and fired in a repeating sequence. Hence, the pressure at the outlet of each combustion tube varies cyclically. A nozzle of the present invention channels the expansion of the pulsed combustion gases from the multiple combustion tubes into a common exhaust stream, in such a manner as to enhance performance in two ways: (1) It reduces the cyclic variations of pressure at the outlets of the combustion tubes so as to keep the pressure approximately constant near the optimum level needed for filling the tubes, regardless of atmospheric pressure at the altitude of operation; and (2) It maximizes the transfer of momentum from the exhaust gas to the engine, thereby maximizing thrust. The figure depicts a typical engine equipped with a nozzle according to the invention. The nozzle includes an interface section comprising multiple intake ports that couple the outlets of the combustion tubes to a common plenum. Proceeding from its upstream to its downstream end, the interface section tapers to a larger cross-sectional area for flow. This taper fosters expansion of the exhaust gases flowing from the outlets of the combustion tubes and contributes to the desired equalization of exhaust combustion pressure. The cross-sectional area for flow in the common plenum is greater than, or at least equal to, the combined cross-sectional flow areas of the combustor tubes. In the common plenum, the exhaust streams from the individual combustion tubes mix to form a single compound subsonic exhaust stream. Downstream of the common plenum is the throat that tapers to a smaller flow cross section. In this throat, the exhaust gases become compressed to form a compound sonic gas stream. Downstream of the throat is an expansion section, which typically has a bell or a conical shape. (The expansion section can be truncated or even eliminated in the case of an air-breathing engine.) After entering the expansion section, the exhaust gases expand rapidly from compound sonic to compound supersonic speeds and are then vented to the environment. The basic invention admits of numerous variations. For example, the combustion tubes can be arranged around the central axin a symmetrical or asymmetrical pattern other than the one shown in the figure. For another example, the flow cross-sectional area(s) of one or more of the intake ports in the interface section, of the common plenum, the throat, and/or the expansion section can be varied, either symmetrically or asymmetrically, to adjust dynamics of the exhaust stream or to direct the thrust vector away from the central axis.

Bratkovich, Thomas E.↗
NASA NTRS1974

Flowfield analysis for successive oblique shock wave-turbulent boundary-layer interactions

Description of a procedure for predicting the flowfields that develop when successive interactions between oblique shock waves and a turbulent boundary layer occur. Such interactions may occur, for example, in engine inlets for supersonic aircraft. A control volume analysis is used to predict changes in the flow field across the interactions. Two bleed flow models are considered. The predicted results are compared with measured results and are shown to be in good agreement when the bleed flow rate is low, or when there is no bleed. Shortcomings of the bleed flow models at higher bleed flow rates are discussed.

Sun, C.-C.↗
DOE OSTI2020

Heat Loss Correction Factor for Fireball Yield Measurements

Fireball yield calculations are performed using Taylor’s radius equation. However, at early times, the x-ray diffusion process has a significant effect on the size of the fireball, and at late times the shock wave degenerates into an acoustic wave. Thus, earlier research determined that the range of validity for Taylor’s equation was in the scaled time regime of 0.004–0.008 s. So, when performing a yield calculation, the only values that are analyzed are those that lie within that range of validity. While Taylor’s equations offer an approximation of the scaled yield, they also rely on several assumptions. For example, the equations assume that the shock wave is adiabatic, when, in fact, the first light pulse on the nuclear test films, which occurs as the shock wave is forming, suggests that the shock wave is not adiabatic. Furthermore, the amount of heat loss from the shock wave that occurs prior to the scaled time regime of 0.004–0.008 s can change significantly as a function of air density. This may cause the shock wave radius to be smaller relative to the assumed adiabatic shock wave. Therefore, the objective of this project was to determine the correction factor for heat transfer, which should be applied in order to produce a more accurate approximation of the weapon’s yield.

45 MILITARY TECHNOLOGY, WEAPONRY, AND NATIONAL DEF↗
NASA NTRS1990

Canard-wing vortex interactions at subsonic through supersonic speeds

The NASA-Ames 6 x 6-foot Transonic/Supersonic Wind Tunnel has been used to conduct a study of canard-wing flowfield interactions at sub-, trans-, and supersonic speeds, giving attention to vortex interactions, vortex breakdown, shock-wave development, and vortex-shock interactions. The results obtained show that the canard-wing flowfield interaction delays vortex breakdown to a higher angle-of-attack at sub- and transonic speeds; while the flowfield interference eliminates shock-induced secondary boundary layer separation on the wing, it does not alter the location and development of a rear shock wave extending laterally across the wing. A canard-induced upwash field accelerates the upward migration of the wing vortex at sub-through-supersonic speeds, but is most pronounced at transonic speeds due to the interaction of the vortical flow with the rear shock wave.

Erickson, Gary E.↗