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In addressing the computational challenges pervasive in engineering where time and cost limitations are key concerns, particularly within the Computational Fluid Dynamics (CFD) domain, Reduced Order Models (ROMs) have emerged as instrumental tools. Focused on reducing computational complexity without intrusively modifying the computational model, this study centres on the strategic application of aerodynamic ROMs, which provide efficient computation of distributed quantities and aerodynamic forces. This work presents ROMs for non-linear aerodynamic applications, integrating principal component analysis (PCA) with Global Local Neural Networks (GLNN). The effectiveness of the proposed methodology is demonstrated by leveraging dependency on the parameter space created with non-linear high-fidelity CFD data, incorporating viscous simulation for a comprehensive approach. Results are first presented for a two-dimensional airfoil case and then for a three-dimensional test case featuring a transonic wing-body-tail transport aircraft configuration (NASA Common Research Model). In transonic flows, the proposed ROMs demonstrate the ability to accurately capture both the location and strength of shocks, as well as forces and moments for unseen prediction points. This highlights the efficiency of the proposed method in navigating complex aerodynamic scenarios, achieving comparable accuracy to full-order modelling but at orders of magnitude less computational time, for unseen parameters outside the ROM training set within the parameter space. • The Proposed ROM reduces the computational time for compressible aerodynamic simulations by orders of magnitude. • It accurately predicts nonlinear aerodynamic load distribution. • It provides predictions for unseen parameters within the parameter space, which were not part of the training data. [ABSTRACT FROM AUTHOR]

Abstract For a high-load low-reaction compressor transonic rotor based on an increase in outlet axial velocity, the stage reaction is decreased by a substantial meridian contraction of the endwall. The axisymmetric endwall contouring has a deep impact on the aerodynamic performance of the rotor. Therefore, nine combinations of endwall contours with three different shapes including a linear wall, a sinusoidal wall and a symmetrical wall of the sinusoidal wall about the linear wall were firstly studied through 3D steady Reynolds Average Navier–Stokes simulations. The results show that the cases with a linear hub have the highest peak pressure ratio and the largest peak efficiency, whereas the cases, featured with a sinusoidal hub, have the widest stall margin, the largest mass flow and the lowest pressure ratio. Endwall contouring not only influences the shock structure across the flow passage but also dominates the location and intensity of separation. In addition, the effect of endwall contouring on loading distribution is also presented and the case with a sinusoidal hub is characterized by aft-loading and the other two cases have a feature of fore-loading. [ABSTRACT FROM AUTHOR]

We present an optimization platform for turbomachinery with complex mesh configuration in a parallel computation environment. A continuous adjoint solver for 3-D viscous internal flow is coded under the same parallel framework as the flow solver. To meet the various permitted extents of reshaping on blade surface and to cut down the computational cost in grid perturbation, a localized two-level mesh deformation method is developed based on Gaussian radial basis function (RBF). This method works efficiently for both the O mesh surrounding the blade and the O–H mesh inside tip gap. In the optimization of the transonic NASA Rotor 67 for high adiabatic efficiency with a mass flow rate constraint, an adjoint sensitivity analysis is conducted. The relations between the design sensitive regions and physical phenomena in internal flow are discussed. Flow fields before and after the adjoint optimization are investigated, including shock system, tip leakage flow, and flow separation. [ABSTRACT FROM AUTHOR]

An adjoint-response surface method is developed to provide efficient surrogate model in a parametrized design space for aerodynamic optimization of turbomachinery blades. Our goal is to improve the adiabatic efficiency or equivalently reduce the entropy generation through blade row with a mass flow rate constraint. Firstly, an aerodynamic sensitivity analysis is conducted with a viscous adjoint method to find the suitable number of control points on the suction surface of the transonic NASA rotor 67. Then quadratic polynomial (QP) based response surfaces of 4, 6 and 8 parameters are examined to validate the advantages of the gradient-enhanced model. In the following 24-parameter aerodynamic design optimization case, a steepest descent optimization (SDO) based on adjoint gradient is conducted, then QP based response surface model is constructed using both the values of cost function and its adjoint gradients with respect to geometry control parameters. We present the geometric features, overall aerodynamic improvements and flow details of optimal designs given by SDO and gradient-enhanced response surface model (GERSM). The effects of blade reshaping on shock system, tip clearance flow and flow separation at hub are examined. Also, off-design performances are analyzed regarding both adiabatic efficiency and stall margin. [ABSTRACT FROM AUTHOR]

Development of casing treatment geometries which will enhance the stable operating envelope of the compressor stage without sacrificing the efficiency is a great challenge for the designers. Bend skewed casing treatment with relatively lower porosities in the presence of plenum chamber is proven to be capable of improving the stall margin and gain in the compressor stage efficiency compared to other conventional casing treatments. Experimental investigations presented in this paper show very encouraging results with substantial improvement in the stable margin of the compressor stage with increase in the stage efficiency. The three bend skewed casing treatments designed for the experimental evaluations have porosities of 21%, 33% and 45%. The steady and transient performances of these three casing treatment geometries are evaluated at two axial extensions of 20% and 40%. Flow survey at the rotor exit shows distinct behaviors of the aerodynamic parameters in the tip region at near stall flow condition. The unsteady velocity measurements are performed with hot wire probe in the tip region ahead of the rotor leading edge to capture the fluctuations in the inlet axial velocity and the results are presented at the stall flow condition. The presence of the casing treatment changes the stalling behavior of the compressor stage. Highest porosity casing treatment results in 40.62% improvement in the stall margin with 0.88% improvement in the compressor stage efficiency at 20% axial extension. [ABSTRACT FROM AUTHOR]

Operational Modal Analysis (OMA) is a promising candidate for flutter testing and Structural Health Monitoring (SHM) of aircraft wings that are passively excited by wind loads. However, no studies have been published where OMA is tested in transonic flows, which is the dominant condition for large civil aircraft and is characterized by complex and unique aerodynamic phenomena. We use data from the HIRENASD large-scale wind tunnel experiment to automatically extract modal parameters from an ambiently excited wing operated in the transonic regime using two OMA methods: Stochastic Subspace Identification (SSI) and Frequency Domain Decomposition (FDD). The system response is evaluated based on accelerometer measurements. The excitation is investigated from surface pressure measurements. The forcing function is shown to be non-white, non-stationary and contaminated by narrow-banded transonic disturbances. All these properties violate fundamental OMA assumptions about the forcing function. Despite this, all physical modes in the investigated frequency range were successfully identified, and in addition transonic pressure waves were identified as physical modes as well. The SSI method showed superior identification capabilities for the investigated case. The investigation shows that complex transonic flows can interfere with OMA. This can make existing approaches for modal tracking unsuitable for their application to aircraft wings operated in the transonic flight regime. Approaches to separate the true physical modes from the transonic disturbances are discussed. [ABSTRACT FROM AUTHOR]

Abstract: High-speed flow research has been sponsored and performed at differing levels of effort since the late 1800s. For example, hypersonic research has experienced numerous cycles since the 1950s, with various periods of high research activity, followed by equally long periods of very low activity. This lack of continuity in high-speed flow research has led to a situation where researchers of one “generation” often do not know what the researchers of previous “generations” have done, mainly due to large losses of institutional knowledge in government, industry, and academic organizations. Therefore, a chronically weak area in research papers, reports, and reviews is the complete identification of critical background documents that form the building blocks and intellectual heritage for modern compressible flow research. A method for systematically determining these critical references is presented in the context of its application to high-speed flow using Citation-Assisted Background, which is based on the assumption that many critical documents tend to be highly cited within the literature, although not necessarily recently. While Citation-Assisted Background is a highly systematic approach for identifying critical references, it is not a substitute for the judgement of the researchers, but rather complements their expertise. In this critical review of high-speed compressible flow, important documents have been identified using Citation-Assisted Background, but other documents have been added by the authors to enhance the picture provided by the highly-cited documents. [Copyright &y& Elsevier]

Abstract: The application of an iterative design method allows the aerodynamic design of engine nacelles and wings being part of complex aircraft configurations. The method couples a flow solver for the solution of the Euler/Navier–Stokes equations on unstructured grids with a design algorithm which solves the transonic potential equation for small perturbations via an inverse formulation. The accuracy and robustness of the design method is investigated for redesigns of engine nacelles and wings. Several design cases demonstrate the efficient application of the method for the aerodynamic design of installed engine nacelles and wings under influence of complex aircraft configurations. One application shows that the method also can be used to minimize aerodynamic installation effects resulting from the engine integration process by sole nacelle or wing contour shaping. A general design methodology for engine nacelles is derived which allows to take into account all aerodynamic effects on complex configurations during the design process. [Copyright &y& Elsevier]

The interaction of aerodynamic and structural forces on a 2-D rectangular wing section oscillating in pitch at transonic flow is studied experimentally. The investigations include free and forced oscillations. The free pitching motion is generated by a symmetric elastic suspension on both sides of the rigid wing, whereas the forced oscillations are induced by an electric actuator on one side of the model. The investigations are based on the simultaneously measured surface pressures and aerodynamic forces while the wing section is oscillating in pitch. The wing model is oscillating about mean incidence angles of zero to two degrees at Mach numbers of 0.50 to 0.85, corresponding to Reynolds numbers 1.5 to 2.2 million. The wing has a supercritical BAC 3-11/RES/30/21 airfoil with a relative thickness of 11% [I.R.M. Moir, AGARD-AR-303, DRA, Farnborough, 1994]. Preliminary investigations include measurements of the steady pressure distribution and flow visualisation at transonic Mach numbers for a fixed mounted wing model with the same airfoil. [Copyright &y& Elsevier]

In a numerical simulation the choice of a turbulence model must be a compromise between physical modelling and computational cost. The CIRA RANS flow solver has been applied, by employing a large set of turbulence models, to typical aerodynamic applications for which certified experimental data are available in literature. The transonic flows over an airfoil and a wing placed in a wind tunnel, both characterized by a strong shock-boundary layer interaction with an induced separation, and the high lift flow around a multi component airfoil are taken into consideration. An evaluation, in terms of accuracy and numerical behaviour, of some common turbulence models ranging from one-equation to high order, using the same code and numerics, is presented.Satisfactory and consistent results have been achieved; the more sophisticated the turbulence model the more accurate the simulation is. The SST Menter κ–ω turbulence model has shown, for the applications investigated, the best compromise between the physical capabilities and the numerical stiffness. [Copyright &y& Elsevier]

Cavity flow studies are generally concerned with observing the effect of geometry changes whilst maintaining a fixed zero angle of attack. Cavities employed as weapons bays will, however, experience a range of angles of attack. This paper presents the first known results showing the effect of flight angle of attack on the aeroacoustic characteristics of an internal weapons bay installed in an uninhabited combat air vehicle (UCAV). The UCAV geometry consisted of a Boeing M219-type cavity in a Boeing UCAV1303 airframe. Numerical simulation was conducted using a full-scale detached eddy simulation model and representative transonic flight conditions. As well as the reference case of zero degrees, data for angles of attack of 3.0, 4.5 and 6.0 degrees were analysed. Experimental data was used to validate the reference computational model, which agreed with the overall fluctuating sound pressure level (OAFPL) to within the experimental uncertainty of 4 dB. Data from the computational model was post-processed with frequency-domain and time-frequency-domain techniques showing that the flow structure within the weapons bay was altered significantly by the angle of attack changes, affecting the mean pressure distribution, frequency spectra and resonant modes. Overall, increasing the angle of attack from 0.0 to 3.0 degrees produced an increment in the acoustic load whilst a further increase tended to affect the resonance mechanism and thereby reduce the coherence and the temporal footprints of the resonant modes. [ABSTRACT FROM AUTHOR]