Transonic Analysis Research Articles

ANALYSIS OF THE BIFURCATING DUCT OF AN INLET PARTICLE SEPARATOR IN TRANSONIC FLOW CONDITIONS

Abstract To increase the reliability of turboprop and turboshaft engines in extreme operating conditions, filtering protections such as Inlet Particle Separators (IPS) can be installed at the intake. The flow inside an IPS is highly 3D and unsteady, with fluctuations especially pronounced when transonic conditions are reached. In this paper, we aim at providing the transonic analysis of an industrial IPS designed for aerodynamic lab testing. As a first step, we define the threshold beyond which sonic conditions are reached in the boundary cross sections of the IPS by means of a 1D semi-empirical model. Second, we select a critical configuration, and we perform CFD simulations to analyze the locations and evolution of the transonic flow inside the IPS. Different models are presented, i.e. steady and unsteady Reynolds-Averaged Navier-Stokes, Detached Eddy Simulation and Large Eddy Simulation, characterized by three levels of resolution. The results show the evolution of some transonic shocks: the steady-state model is only providing information on averaged quantities, while the time-resolved simulations offer a more precise overview in the time domain. The analysis in the frequency domain reveals the frequencies of the transonic fluctuations. Finally, we discuss three alternative designs to effectively improve the operating range and mitigate the risks related to the transonic instabilities, comparing the differences in separation efficiency with respect to the baseline case. The results prove that the optimal design choice is given by the trade-off between operating range, pressure losses and separation efficiency.

High-fidelity aeroelastic transonic analysis using higher-order structural models

A novel computational approach for static aeroelastic analysis of metal and composite wings in transonic flows is proposed. Static aeroelastic analysis is often performed coupling beam/plate structural formulations with low-fidelity inviscid and irrotational aerodynamics. When high subsonic, transonic or supersonic regimes are met, low-fidelity aerodynamics is unable to accurately describe flow separation, viscous phenomena, and/or shock waves, unless suitable corrections are considered. This work combines the use of a variable-order kinematics structural model with Computational Fluid Dynamics (CFD), with the aim of developing a flexible computational aeroelastic framework. In particular, the structural model is based on the Carrera Unified Formulation and Equivalent Plate Modelling, whose governing equations are then solved through the Finite Element Method. The CFD analysis is performed using the high-fidelity open-source software SU2. The fluid–structure interaction is captured resorting to an energetic approach based on the Moving Least Squares patch technique. The generality and flexibility of the developed tool is demonstrated considering: the structural analysis of a wing exhibiting taper ratio, sweep angle, spars and ribs; the aeroelastic static analysis of a transonic AGARD 445.6 wing; and a prototype aeroelastic tailoring study on a composite wing. Comparison with available literature results confirms the robustness of the approach.

고속 병렬처리 기법을 이용한 전기체 항공기 형상의 천음속/초음속 비선형 공탄성 해석

A nonlinear aeroelastic analysis system in transonic and supersonic flows has been developed using high speed parallel processing technique on the network based PC-clustered machines. This paper includes the coupling of advanced numerical techniques such as computational structural dynamics (CSD), finite element method (FEM) and computational fluid dynamics (CFD). The unsteady Euler solver on dynamic unstructured meshes is employed and coupled with computational aeroelastic solvers. Thus it can give very accurate engineering data in the structural and aeroelastic design of flight vehicles. To show the great potential of useful application, transonic and supersonic flutter analyses have been conducted for a complete aircraft model under developing in Korea.

STEADY TWO-DIMENSIONAL TRANSONIC ANALYSIS USING A BOUNDARY INTEGRAL EQUATION METHOD

A boundary integral equation method for the simulation of two-dimensional steady transonic potential flows is presented. The method is based on a conservative differential full-potential formulation. The steady two-dimensional formulation is obtained as the limiting case of the unsteady three-dimensional one, as the iterative method used to obtain steady-state results is a pseudo-time-accurate three-dimensional technique. In the present formulation the full-potential equation appears in the form of a nonlinear wave equation for the velocity potential. All the nonlinear terms, which are expressed in conservative form, are moved to the right-hand side, and treated formally as nonhomogeneous terms. The paper includes a historical review on the development of integral formulations for transonic analysis. Numerical results are obtained for steady two-dimensional transonic flows. Comparisons with existing finite-difference, finite-element, and finite-volume results shows a good agreement. The convergence analysis for an increasing number of grid elements reveals a limit behaviour in good agreement with other numerical methods in both subcritical and supercritical problems. Finally, we present numerical results to demonstrate a remarkable feature of the formulation, that is that the transonic numerical results are quite insensitive to the geometry of the field volume elements; this makes the present formulation particularly appealing for optimal design applications.

ANALYSIS/DESIGN CALCULATIONS OF TRANSONIC FLOWS

SUMMARY Analysis/design calculations of transonic flow are discussed and several improvements are made. The nonisentropic potential method is used to calculate the inviscid transonic flow analysis problem instead of the traditional potential method. An inverse integral 3D boundary layer method is used to calculate the boundary layer in the viscous transonic flow analysis problem. The viscous/inviscid interaction calculations are carried out by a semi-inverse coupling scheme. In design problem calculations, an improved residual-correction method is used. Three individual methods are combined in a global algorithm and computing code. The improvements speed up the convergence, increase applicability and computational efficiency. Some numerical results are given to illustrate that the present method provides an effective engineering tool of high accuracy and efficiency in three dimensional transonic analysis and design situations.

Computation of transonic potential flow past wing-body configurations

A simple approach has been developed to use the two dimensional grid generation method by solving elliptic partial differential equations to generate the three dimensional grid for wingfuselage configurations. This simple method can be applied to generate grids for arbitrary fuselage fitted with any swept wing with dihedral. Three dimensional transonic analysis code TWING, with approximate factorization (AF2) scheme has been suitably used as a flow solver. As an example, RAE-WING-A with body-B2 configuration has been considered. The results obtained have been compared with available numerical and experimental results. It has been observed in the present computations that AF2 scheme is not sensitive to grid stretching.

The computation of transonic analysis and design

In this paper a new high efficiency C-O grid coupled with the potential flow solver based on the finite volume technique is given. It significantly reduces the CPU time and increases the computational efficiency. In order to improve the accuracy of traditional potential method a shock point operator is used to account for entropy correction. Some calculated results of 2D inviscid, viscous/inviscid interaction and 3D inviscid flow indicate that nonisentropic potential method produces results closer to Euler solution as well as experimental data, while its computational efforts are nearly the same as the usual isentropic potential method.

Unsteady transonic algorithm improvements for realistic aircraft applications

Improvements to a time-accurate approximate factorization (AF) algorithm have been implemented for steady and unsteady transonic analysis of realistic aircraft configurations. The algorithm improvements include: an Engquist-Osher (E-O) type-dependent switch to more accurately and efficiently treat regions of supersonic flow, extension of the E-O switch for second-order spatial accuracy in these regions, nonreflecting far field boundary conditions for more accurate unsteady applications, and several modifications which accelerate convergence to steady-state. Calculations are presented for several configurations to evaluate the algorithm modifications. The modifications have significantly improved the stability of the AF algorithm and hence the reliability of the CAP-TSD code in general.

Computational transonic analysis of canted winglets

A computational method developed to provide a transonic analysis for upper/lower surface wing-tip mounted winglets is described Winglets with arbitrary planform, cant and toe angle and airfoil section can be modeled The embedded grid approach provides high flovvfield resolution and the required geometric flexibility In particular coupled Cartesian/cylindrical grid systems are used to model the complex geometry presented by canted upper/lower surface winglets A new rotated difference scheme is introduced in order to maintain the stability of the small disturbance formulation in the presence of large spanvvise velocities Wing and winglet viscous effects are modeled using a two dimensional strip boundary layer analysis. Correlations with wind tunnel and flight test data for three transport configurations are included

A grid overlapping scheme for flowfield computations about multicomponent configurations

An efficient grid interfacing zonal algorithm has been developed for computing the transonic flowfield about three-dimensional multicomponent configurations. The algorithm uses the full-potential formulation and the AF2 fully implicit approximate factorization scheme. The flowfield solution is computed using a component- adaptive grid approach in which separate grids are employed for the individual components in the multicom- ponent configuration, where each component grid is optimized for a particular geometry. The component grids are allowed to overlap, and flowfield information is transmitted from one grid to another through the overlap region. An overlapping grid scheme has been implemented for an isolated wing and a noninteracting wing/pylon/nacelle configuration . Numerical results show that the present algorithm is stable, accurate, and promises to be effective in computing the flowfield about complex configurations. ELIABLE and efficient three-dimensional transonic analysis methods are needed to make realistic and cost- effective predictions of aircraft aerodynamics. Early efforts to predict the transonic flowfield about aircraft multiple- component configurations are based on the transonic small- disturbance formulation.1-3 This allows the geometry of the configuration to be greatly simplified and the surface boundary condition to be applied on a mean approximate surface. Accurate prediction of such flowfields, however, requires the use of the full-potential formulation and the generation of a suitable surf ace-fitted grid. Because each aircraft component (wing, nacelle, fuselage) requires, in general, a grid system that is usually incompatible with the grid systems of the other components, the generation of a single surface-fitted grid for the entire configuration is a difficult task. In such a global grid, control of grid point distribution, skewness, and clustering will be difficult to achieve. Efforts to predict the flowfield about a complete aircraft configuration using a single-grid approach have been made recently by Yu.4 In the present paper the alternate approach of using a component adaptive grid scheme is investigated. The basic idea of this approach is to develop a separate grird for each component of a complex configuration where each individual grid is optimized for a particular component. The flow solver for each component employs the AF2 fully implicit ap- proximate factorization scheme.5 Earlier studies in two dimensions showed that by allowing the component grids to overlap, fast convergence can be achieved.6 This approach has been generalized herein to three dimensions, where an overlapped grid scheme has been im- plemented for an isolated wing and a wing/pylon/nac elle configuration. The application of an embedded grid scheme to an isolated wing geometry was performed primarily as a proof-of- concept study to verify the grid interfacing logic in three dimensions and to determine if single-grid results could be replicated. The application to the wing/pylon/nacelle con- figuration represents a more realistic use of the grid em- bedding technique. The full interacting algorithm has been

Computational aerodynamics on large computers

Three examples of advances in computational aerodynamics—three-dimensional inviscid transonic analysis, design calculations for wings, and the computation of viscous-induced aileron buzz—are reviewed. The importance of large computers in promoting these kinds of advances and the need for continuing increases in computer size and speed are emphasized. NASA's plans for a numerical aerodynamic simulation facility, intended to provide the aeronautics community with a substantial increase in computer capability for the 1980s, are briefly described.

Re´sume´ of Compressor Research at the NACA Langley Laboratory

This paper describes the compressor research program conducted at the NACA Langley Laboratory and highlights the significant developments obtained prior to the discontinuance of this program in 1957. Results of low speed, high speed, and transonic cascade testing are noted. Transonic and supersonic compressor test results and analyses are presented.