Hypersonic Wind Tunnel Nozzle Research Articles

Geometry Optimizatin of Hypersonic Wind Tunnel Nozzle

This paper describes the application of the optimization procedure for a supersonic wind tunnel two-dimensional variable nozzle contour developed in the previous report to the optimization of a hypersonic wind tunnel axisymmetric nozzle contour of Mach number higher than 5.0. In this procedure, several positions are set on the nozzle wall, and using the linearity of a displacement of each position to the flow pattern, the optimum nozzle contour is determined in a few cases of CFD analyses. To take an axisymmetric nozzle of Mach number 10.0 as an example, optimization was carried out. Mach number deviation in the uniform flow region reduced from 1.48% to 0.40% by the optimization.

A genetic algorithm model for high heat flux flow boiling

A new genetic algorithm model is introduced in a recently developed turbulent-boundary-layer scheme for the calculation of heat transfer in high heat flux subcooled boiling flows. Such flows, often desired for cooling of rocket nozzles and nuclear components, are characterized by high fluid velocities and extremely small bubbles that exist in a thin layer adjacent to the heated wall. The current model is based on the solution of the momentum and energy equations with a turbulent diffusivity term based on bubble collapse and associated flow agitation. An important parameter arising in the analysis is the boiling intermittency factor, and the innovation in the present paper is the prediction of this factor through use of a genetic algorithm. The algorithm is employed to simulate the adaptation of a bubble population to the environment in a thermally developing flow. Fourier and Biot numbers for this problem are shown to introduce material and geometric coupling that make the analysis “semiconjugate” in nature. The model is used to predict the heat transfer for a high heat flux water flow typical of that found in the cooling passages of a hypersonic wind tunnel nozzle throat and is compared with some recent high heat flux experimental data.

Compressibility and pressure-gradient correction assessment for turbulent hypersonic flows

NOWLEDGE of the flow development in hypersonic windtunnel nozzles is of primary importance to the design of such nozzles as well as to the characterization of the flowfield in the test section. In the perfect-gas regime, reliable numerical methods have been developed for both of the design work and the flow analysis, but there still exist some deficiencies in predicting the turbulent hypersonic nozzle wall boundary layers. This problem is believed to be a consequence of the deficiencies in the compressible eddy-viscosity turbulence models currently adopted in engineering applications. Specifically, to improve the accuracy of such calculations, the importance of compressibility and pressure-gradient effects should be investigated, since the hypersonic wind-tunnel nozzle flow is characterized by the existence of an increasing main flow Mach number and a varying pressure gradient in the streamwise direction. A numerical study of that nature is presented in this work. Several explicit compressibility and pressure-gradient corrections are introduced to two widely used eddy-viscosity turbulence models for the calculations of hypersonic wind-tunnel nozzle flows. The results are compared with those obtained with the standard models to evaluate the effects of the corrections, and are compared with measurements to verify the accuracy of the calculations and to see the improvements gained. Numerical Modeling The study was carried out by solving three-dimensional Reynolds-averaged Navier-Stokes equations. They are the transport equations to describe the conservation of mass, momentum, and energy with the assumption of a perfect gas. When two-equation turbulence models are employed, the system has two more equations for the turbulent kinetic energy k and its dissipation rate €. In the present work the turbulence is closed by four alternative eddy-viscosity models, among which the Baldwin-Lomax algebraic model1 and the k-€ model of Chien 2 are chosen as the baseline models for the purpose of evaluating the compressibility and pressure-gradient corrections. The other two are their modified versions, respectively, with the modifications outlined below.

Turbulent flow in supersonic and hypersonic nozzles

The calculated turbulent flow properties in a supersonic or a hypersonic wind-tunnel nozzle are affected by the location of laminar-turbulent transition point and the behavior of the particular turbulence models considered. These effects are studied numerically by solving Reynolds-averaged Navier-Stokes equations for two nozzles. Results are reported using either an algebraic turbulence model or two alternative k-e models in which one contains two kinds of explicit compressibility corrections. The numerical method is based on a finite volume technique using a space centered Runge-Kutta scheme which is explicit to the mean flow equations and semi-implicit to the k and e equations. These calculations provide us a better understanding of the mechanism controlling this kind of flowfield. It is found that a small variation of transition point in near throat region has substantial global effects on the computed results downstream. Including explicit compressibility corrections in modeling the hypersonic nozzle flow introduces nonnegligible effects into the solution and improves its accuracy. The turbulence transport effects are also demonstrated.

Flow quality of hypersonic wind-tunnel nozzles designed using computational fluid dynamics

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Aerodynamic design of axisymmetric hypersonic wind-tunnel nozzles using a least-squares/parabolized Navier-Stokes procedure

A new procedure unifying the best of present classical design practices, CFD and optimization procedures, is demonstrated for designing the aerodynamic lines of hypersonic wind tunnel nozzles. This procedure can be employed to design hypersonic wind tunnel nozzles with thick boundary layers where the classical design procedure has been demonstrated to break down. Advantages of this procedure allow full utilization of powerful CFD codes in the design process, solves an optimization problem to determine the new contour, may be used to design new nozzles or improve sections of existing nozzles, and automatically compensates the nozzle contour for viscous effects as part of the unified design procedure.

Aerodynamic design of axisymmetric hypersonic wind-tunnel nozzles

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Atomic recombination in a hypersonic wind-tunnel nozzle

The flow of an ideal dissociating gas through a nearly conical nozzle is considered. The equations of one-dimensional motion are solved numerically assuming a simple rate equation together with a number of different values for the rate constant. These calculations suggest that deviations from chemical equilibrium will occur in the nozzle if the rate constant lies within a very wide range of values, and that, once such a deviation has begun, the gas will very rapidly ’freeze’. The dissociation fraction will then remain almost constant if the flow is expanded further, or even if it passes through a constant area section. An approximate method of solution, making use of this property of sudden ’freezing’ of the flow, has been developed and applied to the problem of estimating the deviations from equilibrium under a wide range of conditions. If all the assumptions made in this paper are accepted, then lack of chemical equilibrium may be expected in the working sections of hypersonic wind tunnels and shock tubes. The shape of an optimum nozzle is derived in order to minimize this departure from equilibrium.It is shown that, while the test section conditions are greatly affected by ’freezing’, the flow behind a normal shock wave is only changed slightly. The heat transfer rate and drag of a blunt body are estimated to be reduced by only about 25% even if complete freezing occurs. However, the shock wave shape is shown to be rather more sensitive to departures from equilibrium.