National Aerospace Laboratories

About: National Aerospace Laboratories is a facility organization based out in Bengaluru, India. It is known for research contribution in the topics: Coating & Corrosion. The organization has 1838 authors who have published 2349 publications receiving 36888 citations.

Electrochemical behavior of superhard nanocomposite coatings of TiN/Si3N4 prepared by reactive DC unbalanced magnetron sputtering

Abstract: Nanocomposite coatings of TiN/Si3N4 are new generation superhard coatings having hardness greater than 40 GPa. Apart from high hardness, these coatings exhibit outstanding properties such as improved wear resistance and high oxidation resistance. These coatings also retain their hardness even after heating in air at higher temperatures. The present study reports electrochemical behavior of superhard nanocomposite coatings of TiN/Si3N4 in 3.5% NaCl solution. The nanocomposite coatings with varying silicon contents have been prepared on steel substrates by reactive direct current unbalanced magnetron sputtering. Single-component TiN and Si3N4 coatings have also been included in the investigation. The electrochemical measurements were performed by potentiodynamic polarization and electrochemical impedance spectroscopy (EIS). The corroded samples were characterized using scanning electron microscopy and energy dispersive X-ray analysis. The nanocomposite coatings exhibited better corrosion resistance than single-phase TiN coatings. Nyquist and Bode plots obtained from the EIS measurements for bare steel substrate, single-component TiN and nanocomposite coatings were used for evaluating the corrosion behavior. The charge transfer resistance and the double layer capacitance were also calculated using EIS data.

On Stokes flow in a semi-infinite wedge

Abstract: Consider Stokes flow in the semi-infinite wedge bounded by the sidewalls ϕ = ±α and the endwall z = 0 Viscous fluid fills the region 0 r < ∞, 0 z < ∞ bounded by these planes; the motion of the fluid is driven by boundary data given on the endwall z = 0 A consequence of the linearity of the problem is that one can treat the velocity field q ( r , ϕ, z ) as the sum of a field q a ( r , ϕ, z ) antisymmetric in ϕ and one symmetric in it, q s ( r , ϕ, z ) It is shown in each of these cases that there exists a real vector eigenfunction sequence v n ( r , ϕ, z ) and a complex vector eigenfunction sequence u n ( r , ϕ, z ), each member of which satisfies the sidewall no-slip condition and has a z -behaviour of the form e − kz It is then shown that one can, for each case, write down a formal representation for the velocity field as an infinite integral over k of the sums of the real and complex eigenfunctions, each multiplied by unknown real and complex scalar functions b n ( k ) and a n ( k ), respectively, that have to be determined from the endwall boundary conditions A method of doing this using Laguerre functions and least squares is developed Flow fields deduced by this method for given boundary data show interesting vortical structures Assuming that the set of eigenfunctions is complete and that the relevant series are convergent and that they converge to the boundary data, it is shown that, in general, there is an infinite sequence of corner eddies in the neighbourhood of the edge r = 0 in the antisymmetric case but not in the symmetric case The same conclusion was reached earlier for the infinite wedge by Sano & Hasimoto (1980) and Moffatt & Mak (1999) A difficulty in the symmetric case when 2α = π/2, caused by the merger of two real eigenfunctions, has yet to be resolved

Capillary damping of inviscid surface waves in a circular cylinder

Abstract: We consider the effect of a wetting condition at the moving contact line on the frequency and damping of surface waves on an inviscid liquid in a circular cylinder. The velocity potential x3C6; and the free surface elevation x3B7; are sought as complex eigenfunction expansions. The x3C6; eigenvalues are the classical ones whereas the x3B7; eigenvalues are unknown and have to be computed so as to satisfy the wetting condition on the contact line and the other free surface conditions x2013; these turn out to be complex in general. A projection of the latter conditions on to an appropriate basis leads to an eigenvalue problem, for the complex frequency x3A9;, which has to be solved iteratively with the wetting condition. The variation of x3A9; with liquid depth h, Bond number Bo, capillary coefficient x3BB; and static contact angle x3B8;c0 is explored for the (1, 0), (2, 0), (0, 1), (3, 0) and (4, 0) modes. The damping vanishes for x3BB; = 0 (pinned-end edge condition) and x3BB;=x221E; (free-end edge condition) with a maximum in the interior while the frequency decreases with increasing x3BB;, approaching limiting values at the endpoints. A comparison with the analytic results of Miles (J. Fluid Mech., vol. 222, 1991, p. 197) for the no-meniscus case and the experimental results of Cocciaro, Faetti, amp; Festa (J. Fluid Mech., vol. 246, 1993, p. 43), where a meniscus is present, is good. The study provides a simple procedure for calculating the inviscid capillary damping associated with the moving contact line in a circular cylinder of finite depth with meniscus effects also being considered.

Visualization of the formation of separation bubbles on a bell-shaped nozzle surface in relation to serious side-load

Abstract: Elimination of serious side-loads are essential in the design of first stage-rocket nozzles. Recent studies have indicated that the major origin of side-loads is a change of separation pattern. In the present paper, shear sensitive liquid crystal (SSLC) was applied to the nozzle wall to visualize changes of separation pattern and to determine the mechanism that produces side-loads. The results showed that SSLC applied to the nozzle wall successfully visualized separation, reattachment and re-separation of the boundary layer and a rapid movement of separation points. The CFD code based on the axisymmetric Navier-Stokes equation was firmly anchored by visualized shear stress distributions, wall pressure distributions and shadow graphs. Introduction Elimination of excessive side-loads during start-up and shut-down transients is one of the most difficult issues in the design of first-stage rocket nozzles. This has long been the subject of investigation but detailed studies on such things as proper contours of the nozzle which would minimize side-loads by optimizing nozzle efficiency have not been conducted. Recently, intensive studies have been conducted to determine the origin of side-loads in the Vulcain nozzle during start-up and shut-down transients [l]. M. Frev et al. |2-4l investigated truncated perfect nozzles and thrust optimized nozzles and found * Researcher, Kakuda Space Propulsion Laboratory, NAL, Member AIAA t Senior researcher, Kakuda Space Propulsion Laboratory, NAL, Member AIAA $ Group leader, Kakuda Space Propulsion Laboratory, NAL, Member AIAA 11 Associate Senior engineer, NASDA Copyright ©2001 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. that the highest side-loads occur in the thrust-optimized nozzle, when the type of separation changes from free shock separation (FSS) to restricted shock separation (RSS), or vice versa. The side-loads of the truncated perfect nozzle, in which only free shock separation occurs, are significantly lower, namely, only about one-third as high [4]. The authors are investigating methodology for the design of the nozzle contour of compressed truncated perfect (CTP) nozzles, proposed by J. D. Hoffman [5], which would optimize nozzle efficiency without severe side-loads. A CTP nozzle contour is obtained by linearly compressing the truncated perfect nozzle (TP nozzle) contour in the axial direction to obtain the desired nozzle length. A discontinuity in the nozzle slope produced in the above compression procedure is e l iminated by introducing a cubic equation which smoothly connects the linearly compressed curve to the initial circular curve. If strong compression is applied to a TP nozzle with a nozzle length longer than that of the original TP nozzle, the nozzle exit angle will be smaller compared with that of the original TP nozzle, resulting in reduction of divergence loss with a probable consequent Fig. 1 Cold flow test facility

Flight control system of Hypersonic Flight Experiment vehicle

Abstract: This paper describes the flight control system of the Hypersonic Flight Experiment vehicle. The experiment was conducted in early 1996. This paper summarizes the vehicle, sensor, on-board computer, control effectors, navigation, guidance, and control laws. In addition, the actual flight is compared with a simulated flight. The obtained flight data show that the flight control system provided enough stability and control performance to achieve the guidance commands. Some differences are observed between the actual and simulated flight, but they are not serious anomalies. (Author)