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Structural patterns of ion and electron currents observed on the wake axes of the Ariel I and the Gemini 10 space vehicles are reexamined, together with relevant theoretical and laboratory 'simulation' studies. Some insight into existing in situ data is provided. The possibility that 'converging-stream' models describe structural features of current enhancements in the wake region of large spacecraft is discussed.
Nonsteady features of a Whitcomb-type supercritical airfoil flowfield were studied in a series of transonic wind tunnel experiments. Data consisted of mean and fluctuating pressures on the airfoil and in the free stream, velocity fluctuations in the wake region, and airfoil vibrations. Fluctuation data were analyzed statistically for intensity, frequency content, and spatial coherence; variations in these parameters were correlated with changes in the mean airfoil flowfield. Surface pressure fluctuation intensity was influenced primarily by the location and motion of the upper-surface shock, the existence of separation, and (downstream of the shock) the development and upstream propagation of 'shocklets'. Power spectra of pressure fluctuations showed characteristics differences up- and downstream of the shock. Chordwise and spanwise coherences were considerably reduced in the presence of separation. Downwash fluctuations at the edge of the wake were noticeably affected only by the development of trailing-edge flow separation.
Some aspects of sound generation by the interaction between a wedge and the wake of a circular cylinder in a uniform incompressible stream are studied experimentally. There is a minimum separation distance between the cylinder and the wedge when a tone is first generated. As the separation distance increases beyond the minimum, the frequency increases and asymptotically approaches the frequency normally present in the free wake. The Strouhal number defined as the ratio fd/U where f is the frequency, d is the cylinder diameter and U the free-stream speed is found to be in general a function of the Reynolds number (based on d) and h/d, where h is the separation distance. The minimum distance is in general a function of the Reynolds number. Flow visualization studies show that considerable differences exist between the flow below the minimum separation distance and just above it.
The paper reviews NASA's efforts on wake vortex turbulence minimization by aerodynamic design or retrofit modifications to large transport aircraft. Theoretical and experimental (ground-based and flight) results are presented which show that the adverse effects of a vortex wake produced by a large aircraft on a small following aircraft can be reduced significantly.
A form of the wall-wake profile, which is applicable to flows with heat transfer, and for which a variation in y = O at y = delta, was suggested. The modified profile, which takes into account the effect of turbulent Prandtl number, was found to provide a good representation of experimental data for a wide range numbers and heat transfer. The Cf values which are determined by a least squares fit of the profile to the data agree well with values which were measured by the floating element technique. In addition, the values of delta determined by the fit correspond more closely to the outer edge of the viscous flow region than those obtained with earlier versions of the wall-wake profile.
The effect of wing wake on the lift of the horizontal tail surfaces was studied. In the development of expressions for this effect, the growth of wing circulation and wing wake, the time interval represented by the tail length, and the development of lift by the tail were considered. The theory has been applied to a specific case to show the magnitude of the effect to be expected. It is shown that, for motions below a certain frequency, the development of lift by the tail may be represented by a simple lag function. The lag is, however, somewhat greater than that indicated by the tail length.
Some recent research relating to the nature of the lift-induced vortex wakes behind large aircraft was reviewed and the scaling laws that permit a comparison of results from ground facilities with those from flight test were provided. The maximum rotational velocities in the wake are shown to depend on a span loading shape parameter and on a characteristic length of persistence behind the aircraft. The effects of Reynolds number are also shown.
A new model for the near-wake studies is presented. It pertains to a tenuous plasma medium with negligible magnetic field effects. The wake-generating body moves at a mesothermal speed, i.e., intermediate between the thermal speeds of ions and electrons of the ambient plasma which is bithermal (Ti less than Te) and has a Debye length much smaller than the body size. The model is constructed by using the concept that the cavity left by a rapidly moving body acts as a generator of finite-amplitude ion acoustic waves which are excited by an initial discontinuity of plasma density at the leading edge of the cavity. This ionic wave motion, represented by fluid equations, is represented approximately by the time dependent one-dimensional Korteweg-de Vries equation. Numerical and analytic solutions are both given.
The paper suggests a new form of the wall-wake profile which is applicable to flows with heat transfer, and for which the velocity gradient at the boundary-layer edge is zero. The modified profile - which takes into account the effect of turbulent Prandtl number - is found to provide a good representation of experimental data for a wide range of Mach numbers and heat transfer. The skin friction coefficient values determined by a least squares fit of the profile to the data are in good agreement with values measured by the floating element technique. The values of boundary-layer thickness determined by the fit correspond more closely to the outer edge of the viscous flow region than those obtained with earlier versions of the wall-wake profile.
Piloted six-degrees-of-freedom motion simulator investigations were conducted at the NASA Ames Research Center to determine criteria relating the hazard posed by a wake vortex encounter to the response of the encountering airplane. These investigations demonstrated that wake vortex encounters can be realistically reproduced on a simulator, established that the maximum bank angle due to the encounter provides the best correlation with the pilot's subjective assessment of the hazard, and determined hazard boundaries in terms of maximum bank angle for two classes of jet transport aircraft.
The several wake vortices which originate from aircraft wingtips, flap edges, engine pylons, etc. usually merge, in the far field, to form a single pair whose structure determines the hazard posed to encountering aircraft. To gain an understanding of the process whereby vortices merge and disperse, a numerical study was made of the interaction of two-dimensional, time-dependent, inviscid vortical regions. It was found that discrete boundaries, which depend on the structure and spacing of the vortices, distinguish merging from nonmerging situations. Furthermore, certain arrays of finite vortex cores that alternate in sign were found to undergo division and merging that may be useful in alleviating the hazard posed by aircraft wakes.
The generation of blade passing frequency noise due to rotor wakes interaction with the downstream stator is examined using a recent model for the unsteady aerodynamic forces developed by Goldstein and Atassi. The effects of blade geometry and loading are investigated. It is shown that at low reduced frequencies loading significantly reduces the noise level while at high frequencies loading substantially increases the sound level radiated. It is also shown that for the same loading a higher angle of attack generates more noise particularly at higher reduced frequencies. The noise level shows a quasi-periodic pattern with pronounced minima and maxima as the frequency increases. These conclusions are valid for different wake models.
A laser velocimeter capable of rapidly scanning a flow field while simultaneously sensing two components of the velocity was used to measure the vertical and streamwise velocity structure 1.5 spans downstream in the wake of a model typical of a large subsonic transport (Boeing 747). This flow field was modeled by a superposition of axisymmetric vortices with finite cores. This theoretical model was found to agree with the measured velocities everywhere except where two vortices were in close proximity. Vortex strengths derived from the span loading on the wing as predicted by vortex-lattice theory also agree with the present measurements. The axisymmetric vortex model used herein is a useful tool for analytically investigating the vortex wakes of aircraft.
The vacuum environment in the wake region of presently planned large space vehicles is calculated using simplified models of the particle fluxes from the various sources. The fluxes which are calculated come directly from the ambient, are due to ambient particles backscattered from spacecraft emissions, and are due to self scattering of spacecraft emissions. Using nominal values for the surface emissions, the flux density environment behind a large unmanned craft at 550 km altitude is calculated. Calculations indicate that the flux density on a wake vacuum experiment conducted in the vicinity of the shuttle is substantially greater than that behind unmanned craft.
The dynamic interaction of aircraft wake vortices was investigated using both inviscid and viscous models. For the viscous model, a computer code was developed using a second-order closure model of turbulent transport. The phenomenon of vortex merging which results in the rapid aging of a vortex wake was examined in detail. It was shown that the redistribution of vorticity during merging results from both convective and diffusive mechanisms.
A development program has demonstrated that hot-wire anemometry can be used successfully on an aircraft in flight to make measurements of wake vortices produced by another aircraft. The probe, whose wires were made of platinum/rhodium, 10 microns in diameter, provides unambiguous results for inflow angles less than about 35 deg. off the probe axis. The high frequency response capability of the hot-wire system allows detailed measurement of the flow structure, and the study of aircraft hazards associated with wake turbulence.
Horseshoe-like vortices, induced by wakes in the stagnation region of bluff bodies, are proposed as an efficient mechanism for augmentation of convective heat transfer. The vortex 'flow module' induced by single or multiple wakes, which had not been observed previously, was first documented and the resulting flow field was studied using various visualization techniques and hot-wire anemometry. In an attempt to understand the driving force behind this flow module, the conditions at which incipient formation of the vortices occurs were investigated. Existence of such a threshold is essential and was hitherto an open question in analytical studies of stability of flow in stagnation region. Finally, effects of the flow module on heat transfer from a cylinder were measured.
A computational technique applicable to analysis of supersonic transport (SST) wake photochemistry and diffusion is presented. Sensitivity studies of SST effluent effects upon ozone depletion are facilitated by the computational rapidity of the method. The article compares results from other studies and predictions of some variables related to global NOx input. Results indicate that the NO/NO2 ratio in an SST wake at photochemical equilibrium is a sensitive function of photolysis rates.