A Pseudo-Compressibility Method for Solving Inverse Problems based on the 3D Incompressible Euler Equations

A numerical technique to solve the three-dimensional inverse problems that arise in aerodynamic design is presented. The approach, which is well established for compressible flows, is extended to the incompressible case via artificial compressibility preconditioning. The modified system of equations is integrated with a characteristic-based Godunov method. The solution of the inverse problem is given as the steady state of an ideal transient during which the flow field assesses itself to the boundary conditions, which are prescribed as design data, by changing the boundary contour. The main aspects of the Eulerian-Lagrangian numerical procedure are illustrated and the results are validated by comparisons with theoretical solutions and experimental results

Active Flow Control in Supersonic Nozzles

Active flow control in supersonic nozzle mainly concerns Fluidic Thrust Vectoring (FTV) strategies. Thrust Vectoring in fixed symmetric nozzles can be obtained by generating a local perturbation at wall causing flow separations, asymmetric pressure distributions and therefore, the vectoring of the primary jet thrust. The control action can be steady, e.g. continuous blowing, or pulsating, e.g. by synthetic jet actuators. In the paper a numerical procedure is explained, which is able to deal with most of the FTV strategies as well as with continuous and pulsating flow excitation. The flow governing equations are solved according to a finite volume discretization of the compressible URANS equations. The effectiveness of different combinations of FTV strategies and flow actuations are presented. The numerical results obtained are compared with the experimental data found in the open literature

A New Steklov-Poincaré Numerical Technique for Solving Prandtl-Batchelor Flows

The Prandtl-Batchelor flow model is a well-known asymptotic solution of the Navier-Stokes equations often used as a paradigm model of wake past bluff bodies. The main concern is the derivation of vortex equilibria and stability in symmetric and asymmetric configurations. The numerical solution of such class of problems requires an high resolution of the flow in the wake regions and a wide grid to accurately match asymptotic conditions at infinity. In this work a numerical technique based on the Steklov-Poincaré iterative scheme is proposed in order to match the asymptotic conditions and a high resolution on the wake. The case of body endowed with sharp edges is also considered

Numerical Simulation of Fluidic Thrust-Vectoring

The paper focuses on a computational method for the investigation of Fluidic Thrust Vectoring (FTV). Thrust vectoring in symmetric nozzles is obtained by secondary flow injections that cause local flow separations, asymmetric pressure distributions and, therefore, the vectoring of the primary jet thrust. The methodology proposed here can be applied for studying numerically most of the strategies for fluidic thrust vectoring, as shock-vector control, sonic-plane skewing and the counterflow method. The computational technique is based on a well-assessed mathematical model. The flow governing equations are solved according to a finite volume discretization technique of the compressible RANS equations coupled with the Spalart-Allmaras turbulence model. Second order accuracy in space and time is achieved using an Essentially Non Oscillatory scheme. For validation purposes, the proposed numerical tool is used for the simulation of thrust vectoring based on the dual-throat nozzle concept. Nozzle performances and thrust vector angles are computed for a wide range of nozzle pressure ratios and secondary flow injection rates. The numerical results obtained are compared with the experimental data available in the open literature

Optimal Aerodynamic Design of Scramjet Facility Nozzles

An hybrid approach to optimization is applied to the aerodynamic design of a device that acts as a facility nozzle and inlet distortion generator for a direct-connect scramjet combustor test-rig. The solution of the design problem is sought by using different approaches of CFD analysis and optimization tools. The initial design is based on the solution of an inverse problem coupled to optimization by genetic algorithms. The result of the optimization is then tested by CFD simulations of the direct-connect facility and the shock reflection patterns are compared toat of the in-flight configuration. The verification of nozzle design and the assessment of the prescribed flow distortion are then carried out by numerical simulations of the flowfield of the whole direct-connect facility by an URANS solver. The proposed procedure is checked by designing a facility nozzle and distortion generator system for a scramjet model available in the open literature

Studio e Controllo di Strutture Vorticose di Parete

At some flow configurations a strong vortex is standing close to the wall and creates a sliding surface which moves and prevents flow separations. In the literature this phenomenon is known as the "snow-cornice effect" . An analitycal model for steady solutions is proposed here, which is able to predict the main features of the flow control mechanism. The tools of the theory of the Hamiltonian dynamical systems are applied to determine the strength and the the position of the trapped vortex and to discuss its stability. The unsteady evolution of a distributed-vorticity region is then investigated numerically by using a high-order blob vortex method. To speed up the computations, a Fast multipole Method is used, which has been extended to general domains by a conformal mapping technique. The transient from the impulsive start to the final, on average steady, flow characterized by the trapped vortex structure is observed for a flow configuration with a single cavity ( the Ringleb Flow), for flows with periodic cavities, and for flows on rotor with N cusps and cavitie

ACTIVE FLOW CONTROL OF AN OVER-EXPANDED NOZZLE BY SHOCK VECTOR CONTROL

Thrust vectoring obtained by the nozzle flow manipulation technique known as Shock Vector Control (SVC) is investigated numerically. In the shock vector control method, a shock structure is generated in the main flow by using transversal continuous blowing. The pressure distribution on the nozzle walls becomes asymmetric, thus giving rise to a lateral force. The open-loop response of the nozzle and the thrust vectoring effectiveness/controllability are investigated by using a CFD tool based on the compressible URANS equations. Nozzle performances and thrust vector angles have been computed for different nozzle pressure ratios in the range of over-expanded conditions. The latter represent the worst case, where the shock structure generated by the control is amplified by the re-compression requirements imposed by the external ambient pressure

Thrust Vectoring of a Fixed Axisymmetric Supersonic Nozzle Using the Shock-Vector Control Method

The application of the Shock Vector Control (SVC) approach to an axysimmetric supersonic nozzle is studied numerically. SVC is a Fluidic Thrust Vectoring (FTV) strategy that is applied to fixed nozzles in order to realize jet-vectoring effects normally obtained by deflecting movable nozzles. In the SVC method, a secondary air flow injection close to the nozzle exit generates an asymmetry in the wall pressure distribution and side-loads on the nozzle, which are also lateral components of the thrust vector. SVC forcing of the axisymmetric nozzle generates fully three-dimensional flows with very complex structures that interact with the external flow. In the present work, the experimental data on a nozzle designed and tested for a supersonic cruise aircraft are used for validating the numerical tool at different flight Mach numbers and nozzle pressure ratios. Then, an optimal position for the slot is sought and the fully 3D flow at flight Mach number M∞=0.9 is investigated numerically for different values of the SVC forcing

Optimal Inverse Method for Turbomachinery Design

An adjoint optimization method based on the solution of an inverse problem is proposed. In this formulation, the distributed control is a flow variable on the domain boundary, for example pressure. The adjoint formulation delivers the functional gradient with respect to such flow variable distribution, and a descent method can be used for optimization. The flow constraints are easily imposed in the parametrization of the controls, thus those problems with many strict constraints on the flow solution can be solved very efficiently. Conversely, the geometric constraints are imposed either by additional partial differential equations, or by penalization. Constraining the geometric solution, the classical limitations of the inverse problem design are overcome. Two examples pertaining to internal flows are give

Optimal aerodynamic design of hypersonic inlets by using streamline-tracing techniques

Rectangular-to-Ellipse Shape Transition (REST) inlets are a class of inward turning inlets designed for hypersonic flight. The aerodynamic design of REST inlets involves very complex flows and shock-wave patterns. These inlets are used in highly integrated propulsive systems. Often the design of these inlets may require many geometrical constraints at different cross-section. In present work a design approach for hypersonic inward-turning inlets, adapted for REST inlets, is coupled with a multi-objective optimization procedure. The automated procedure iterates on the parametric representation and on the numerical solution of a base flow from which the REST inlet is generated by using streamline tracing and shape transition algorithms. The typical design problem of optimizing the total pressure recovery and mass flow capture of the inlet is solved by the proposed procedure. The accuracy of the optimal solutions found is discussed and the performances of the designed REST inlets are investigated by means of fully 3-D Euler and 3-D RANS analyses